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NASA/SP— 1 998-7037/SUPPL387 
November 13, 1998 



AERONAUTICAL 
ENGINEERING 



A CONTINUING BIBLIOGRAPHY WITH INDEXES 




National Aeronautics and 
Space Administration 
Langley Research Center 

Scientific and Technical 
Information Program Office 



The NASA STI Program Office ... in Profile 



Since its founding, NASA has been dedicated 
to the advancement of aeronautics and space 
science. The NASA Scientific and Technical 
Information (STI) Program Office plays a key 
part in helping NASA maintain this important 
role. 

The NASA STI Program Office is operated by 
Langley Research Center, the lead center for 
NASA's scientific and technical information. 
The NASA STI Program Office provides access 
to the NASA STI Database, the largest collection 
of aeronautical and space science STI in the 
world. The Program Office is also NASA's 
institutional mechanism for disseminating the 
results of its research and development activities. 
These results are published by NASA in the 
NASA STI Report Series, which includes the 
following report types: 

• TECHNICAL PUBLICATION. Reports of 
completed research or a major significant 
phase of research that present the results of 
NASA programs and include extensive data or 
theoretical analysis. Includes compilations of 
significant scientific and technical data and 
information deemed to be of continuing 
reference value. NASA's counterpart of peer- 
reviewed formal professional papers but has 
less stringent limitations on manuscript length 
and extent of graphic presentations. 

• TECHNICAL MEMORANDUM. Scientific 
and technical findings that are preliminary or 
of specialized interest, e.g., quick release 
reports, working papers, and bibliographies 
that contain minimal annotation. Does not 
contain extensive analysis. 

• CONTRACTOR REPORT. Scientific and 
technical findings by NASA-sponsored 
contractors and grantees. 



• CONFERENCE PUBLICATION. Collected 
papers from scientific and technical 
conferences, symposia, seminars, or other 
meetings sponsored or cosponsored by NASA. 

• SPECIAL PUBLICATION. Scientific, 
technical, or historical information from 
NASA programs, projects, and missions, 
often concerned with subjects having 
substantial public interest. 

• TECHNICAL TRANSLATION. 

English- language translations of foreign 
scientific and technical material pertinent to 
NASA's mission. 

Specialized services that complement the STI 
Program Office's diverse offerings include 
creating custom thesauri, building customized 
databases, organizing and publishing research 
results . . . even providing videos. 

For more information about the NASA STI 
Program Office, see the following: 

• Access the NASA STI Program Home Page at 
http:llwww. sti. nasa.gov 

• E-mail your question via the Internet to 
help@sti.nasa.gov 

• Fax your question to the NASA STI Help Desk 
at (301) 621-0134 

• Telephone the NASA STI Help Desk at 
(301) 621-0390 

• Write to: 

NASA STI Help Desk 

NASA Center for AeroSpace Information 

7121 Standard Drive 

Hanover, MD 21076-1320 



Introduction 



This supplemental issue of Aeronautical Engineering, A Continuing Bibliography with Indexes 
(NASA/SP — 1998-7037) lists reports, articles, and other documents recently announced in the 
NASA STI Database. 

The coverage includes documents on the engineering and theoretical aspects of design, construction, 
evaluation, testing, operation, and performance of aircraft (including aircraft engines) and associ- 
ated components, equipment, and systems. It also includes research and development in aerodynam- 
ics, aeronautics, and ground support equipment for aeronautical vehicles. 

Each entry in the publication consists of a standard bibliographic citation accompanied, in most 
cases, by an abstract. 

The NASA CASI price code table, addresses of organizations, and document availability informa- 
tion are included before the abstract section. 

Two indexes — subject and author are included after the abstract section. 



SCAN Goes Electronic! 

If you have electronic mail or if you can access the Internet, you can view biweekly issues of SCAN 
from your desktop absolutely free! 

Electronic SCAN takes advantage of computer technology to inform you of the latest worldwide, 
aerospace-related, scientific and technical information that has been published. 

No more waiting while the paper copy is printed and mailed to you. You can view Electronic SCAN 
the same day it is released — up to 191 topics to browse at your leisure. When you locate a publication 
of interest, you can print the announcement. You can also go back to the Electronic SCAN home page 
and follow the ordering instructions to quickly receive the full document. 

Start your access to Electronic SCAN today. Over 1,000 announcements of new reports, books, con- 
ference proceedings, journal articles. ..and more — available to your computer every two weeks. 

i-j For Internet access to E-SCAN, use any of the 

"J^ffl* J ifitfl^ following addresses: 

* /^/itllP*' -«-i I http://www.sti.nasa.gov 

l7lljE^ • ftp.sti.nasa.gov 

*■ gopher.sti.nasa.gov 

To receive a free subscription, send e-mail for complete information about the service first. Enter 
scan@sti.nasa.gov on the address line. Leave the subject and message areas blank and send. You 
will receive a reply in minutes. 

Then simply determine the SCAN topics you wish to receive and send a second e-mail to 
listserve@sti.nasa.gov. Leave the subject line blank and enter a subscribe command in the message 
area formatted as follows: 

Subscribe <desired list> <Your name> 

For additional information, e-mail a message to help@sti.nasa.gov. 

Phone: (301) 621-0390 

Fax: (301) 621-0134 

Write: NASA STI Help Desk 

NASA Center for AeroSpace Information 
7121 Standard Drive 
Hanover, MD 21076-1320 

Looking just for Aerospace Medicine and Biology reports? 



Although hard copy distribution has been discontinued, 
you can still receive these vital announcements through 
your E-SCAN subscription. Just subscribe SCAN-AEROMED 
in the message area of your e-mail to listserve@sti.nasa.gov. 



Table of Contents 



Records are arranged in categories 1 through 19, the first nine coming from the Aeronautics division 
of STAR, followed by the remaining division titles. Selecting a category will link you to the collection 
of records cited in this issue pertaining to that category. 



Includes aerodynamics of bodies, combinations, wings, rotors, and control surfaces; and 
internal flow in ducts and turbomachinery. 



Includes passenger and cargo air transport operations; and aircraft accidents. 



Includes digital and voice communication with aircraft; air navigation systems (satellite and 
ground based); and air traffic control. 



Includes aircraft simulation technology. 

Includes cockpit and cabin display devices; and flight instruments. 



Includes prime propulsion systems and systems components, e.g., gas turbine engines and 
compressors; and onboard auxiliary power plants for aircraft. 



Includes aircraft handling qualities; piloting; flight controls; and autopilots. 



Includes airports, hangars and runways; aircraft repair and overhaul facilities; wind tunnels; 
shock tubes; and aircraft engine test stands. 



Includes astronautics (general); astrodynamics; ground support systems and facilities 
(space); launch vehicles and space vehicles; space transportation; space communications, 
spacecraft communications, command and tracking; spacecraft design, testing and perfor- 
mance; spacecraft instrumentation; and spacecraft propulsion and power. 



Includes chemistry and materials (general); composite materials; inorganic and physical 
chemistry; metallic materials; nonmetallic materials; propellants and fuels; and materials 
processing. 



Includes engineering (general); communications and radar; electronics and electrical engi- 
neering; fluid mechanics and heat transfer; instrumentation and photography; lasers and 
masers; mechanical engineering; quality assurance and reliability; and structural mechanics. 



Includes geosciences (general); earth resources and remote sensing; energy production and 
conversion; environment pollution; geophysics; meteorology and climatology; and ocean- 
ography. 



Includes life sciences (general); aerospace medicine; behavioral sciences; man/system 
technology and life support; and space biology. 



Includes mathematical and computer sciences (general); computer operations and hardware; 
computer programming and software; computer systems; cybernetics; numerical analysis; 
statistics and probability; systems analysis; and theoretical mathematics. 



Includes physics (general); acoustics; atomic and molecular physics; nuclear and high- 
energy; optics; plasma physics; solid-state physics; and thermodynamics and statistical 
physics. 



Includes social sciences (general); administration and management; documentation and 
information science; economics and cost analysis; law, political science, and space policy; 
and urban technology and transportation. 



Includes space sciences (general); astronomy; astrophysics; lunar and planetary exploration; 
solar physics; and space radiation. 



Two indexes are available. You may use the find command under the tools menu while viewing the 
PDF file for direct match searching on any text string. You may also view the indexes provided, for 
searching on NASA Thesaurus subject terms and author names. 



Selecting an index above will link you to that comprehensive listing. 



Select Availability IflfO for important information about NASA Scientific and Technical Infor- 
mation (STI) Program Office products and services, including registration with the NASA Center 
for AeroSpace Information (CASI) for access to the NASA CASI TRS (Technical Report Server), 
and availability and pricing information for cited documents. 



The New NASA Video 
Catalog is He 







To order your 11 copy, 

call the NASA STI Help Desk at 

(301)621-0390, 

fax to 

(301)621-0134, 

e-mail to 

help@sti.nasa.gov, 

or visit the NASA STI Program 

homepage at 



(Select STI Program Bibliographic Announcements) 



Explore the Universe! 



f 



The mission of the NASA Scientific and Technical (STI) Program Office is to quickly, efficiently, 
and cost-effectively provide the NASA community with desktop access to STI produced by NASA 
and the world's aerospace industry and academia. In addition, we will provide the aerospace 
industry, academia, and the taxpayer access to the intellectual scientific and technical output and 
achievements of NASA. 

Eligibility and Registration for NASA STI Products and Services 

The NASA STI Program offers a wide variety of products and services to achieve its mission. Your 
affiliation with NASA determines the level and type of services provided by the NASA STI 
Program. To assure that appropriate level of services are provided, NASA STI users are requested to 
register at the NASA Center for AeroSpace Information (CASI). Please contact NASA CASI in one 
of the following ways: 

E-mail: help@sti.nasa.gov 

Fax: 301-621-0134 

Phone: 301-621-0390 

Mail: ATTN: Registration Services 

NASA Center for AeroSpace Information 

7121 Standard Drive 

Hanover, MD 21076-1320 

Limited Reproducibility 

In the database citations, a note of limited reproducibility appears if there are factors affecting the 
reproducibility of more than 20 percent of the document. These factors include faint or broken type, 
color photographs, black and white photographs, foldouts, dot matrix print, or some other factor that 
limits the reproducibility of the document. This notation also appears on the microfiche header. 



NASA Patents and Patent Applications 

Patents and patent applications owned by NASA are announced in the STI Database. Printed copies 
of patents (which are not microfiched) are available for purchase from the U.S. Patent and 
Trademark Office. 

When ordering patents, the U.S. Patent Number should be used, and payment must be remitted in 
advance, by money order or check payable to the Commissioner of Patents and Trademarks. Prepaid 
purchase coupons for ordering are also available from the U.S. Patent and Trademark Office. 



NASA patent application specifications are sold in both paper copy and microfiche by the NASA 
Center for AeroSpace Information (CASI). The document ID number should be used in ordering 
either paper copy or microfiche from CASI. 

The patents and patent applications announced in the STI Database are owned by NASA and are 
available for royalty-free licensing. Requests for licensing terms and further information should be 
addressed to: 

National Aeronautics and Space Administration 

Associate General Counsel for Intellectual Property 

Code GP 

Washington, DC 20546-0001 

Sources for Documents 

One or more sources from which a document announced in the STI Database is available to the 
public is ordinarily given on the last line of the citation. The most commonly indicated sources and 
their acronyms or abbreviations are listed below, with an Addresses of Organizations list near the 
back of this section. If the publication is available from a source other than those listed, the publisher 
and his address will be displayed on the availability line or in combination with the corporate source. 

Avail: NASA CASI. Sold by the NASA Center for AeroSpace Information. Prices for hard copy 
(HC) and microfiche (MF) are indicated by a price code following the letters HC or MF in 
the citation. Current values are given in the NASA CASI Price Code Table near the end of 
this section. 

Note on Ordering Documents: When ordering publications from, NASA CASI, use the document ID number 
or other report number. It is also advisable to cite the title and other bibliographic identification. 

Avail: SOD (or GPO). Sold by the Superintendent of Documents, U.S. Government Printing 
Office, in hard copy. 

Avail: BLL (formerly NLL): British Library Lending Division, Boston Spa, Wetherby, Yorkshire, 
England. Photocopies available from this organization at the price shown. (If none is given, 
inquiry should be addressed to the BLL.) 

Avail: DOE Depository Libraries. Organizations in U.S. cities and abroad that maintain 
collections of Department of Energy reports, usually in microfiche form, are listed in 
Energy Research Abstracts. Services available from the DOE and its depositories are 
described in a booklet, DOE Technical Information Center — Its Functions and Services 
(TID-4660), which may be obtained without charge from the DOE Technical Information 
Center. 

Avail: ESDU. Pricing information on specific data, computer programs, and details on ESDU 
International topic categories can be obtained from ESDU International. 

Avail: Fachinformationszentrum Karlsruhe. Gesellschaft fur wissenschaftlich-technische 
Information mbH 76344 Eggenstein-Leopoldshafen, Germany. 



Avail: HMSO. Publications of Her Majesty's Stationery Office are sold in the U.S. by Pendragon 
House, Inc. (PHI), Redwood City, CA. The U.S. price (including a service and mailing 
charge) is given, or a conversion table may be obtained from PHI. 

Avail: Issuing Activity, or Corporate Author, or no indication of availability. Inquiries as to the 
availability of these documents should be addressed to the organization shown in the 
citation as the corporate author of the document. 

Avail: NASA Public Document Rooms. Documents so indicated may be examined at or purchased 
from the National Aeronautics and Space Administration (JBD-4), Public Documents 
Room (Room 1H23), Washington, DC 20546-0001, or public document rooms located at 
NASA installations, and the NASA Pasadena Office at the Jet Propulsion Laboratory. 

Avail: NTIS. Sold by the National Technical Information Service. Initially distributed microfiche 
under the NTIS SRIM (Selected Research in Microfiche) are available. For information 
concerning this service, consult the NTIS Subscription Section, Springfield, VA 22161. 

Avail: Univ. Microfilms. Documents so indicated are dissertations selected from Dissertation 
Abstracts and are sold by University Microfilms as xerographic copy (HC) and microfilm. 
All requests should cite the author and the Order Number as they appear in the citation. 

Avail: US Patent and Trademark Office. Sold by Commissioner of Patents and Trademarks, U.S. 
Patent and Trademark Office, at the standard price of $1.50 each, postage free. 

Avail: (US Sales Only). These foreign documents are available to users within the United States 
from the National Technical Information Service (NTIS). They are available to users 
outside the United States through the International Nuclear Information Service (IMS) 
representative in their country, or by applying directly to the issuing organization. 

Avail: USGS. Originals of many reports from the U.S. Geological Survey, which may contain 
color illustrations, or otherwise may not have the quality of illustrations preserved in the 
microfiche or facsimile reproduction, may be examined by the public at the libraries of the 
USGS field offices whose addresses are listed on the Addresses of Organizations page. The 
libraries may be queried concerning the availability of specific documents and the possible 
utilization of local copying services, such as color reproduction. 



Addresses of Organizations 



British Library Lending Division 
Boston Spa, Wetherby, Yorkshire 
England 

Commissioner of Patents and Trademarks 
U.S. Patent and Trademark Office 
Washington, DC 20231 

Department of Energy 
Technical Information Center 
P.O. Box 62 
Oak Ridge, TN 37830 

European Space Agency- 
Information Retrieval Service ESRIN 
Via Galileo Galilei 
00044 Frascati (Rome) Italy 

ESDU International 
27 Corsham Street 
London 
Nl 6UA 
England 

Fachinformationszentrum Karlsruhe 

Gesellschaft fur wissenschaftlich-technische 
Information mbH 

76344 Eggenstein-Leopoldshafen, Germany 

Her Majesty's Stationery Office 
P.O. Box 569, S.E. 1 
London, England 

NASA Center for AeroSpace Information 
7121 Standard Drive 
Hanover, MD 21076-1320 

(NASA STI Lead Center) 

National Aeronautics and Space Administration 

Scientific and Technical Information Program Office 

Langley Research Center - MS 157 

Hampton, VA 23681 



National Technical Information Service 
5285 Port Royal Road 
Springfield, VA 22161 

Pendragon House, Inc. 
899 Broadway Avenue 
Redwood City, CA 94063 

Superintendent of Documents 
U.S. Government Printing Office 
Washington, DC 20402 

University Microfilms 
A Xerox Company 
300 North Zeeb Road 
Ann Arbor, MI 48106 

University Microfilms, Ltd. 
Tylers Green 
London, England 

U.S. Geological Survey Library National Center 

MS 950 

12201 Sunrise Valley Drive 

Reston, VA 22092 

U.S. Geological Survey Library 
2255 North Gemini Drive 
Flagstaff, AZ 86001 

U.S. Geological Survey 
345 Middlefield Road 
Menlo Park, CA 94025 

U.S. Geological Survey Library 
Box 25046 

Denver Federal Center, MS914 
Denver, CO 80225 



NASA CASI Price Code Table 

(Effective July 1,1998) 



U.S., Canada, 




Code & Mexico 


Foreigi 


AOl $ 8.00 


. . $ 16.00 


A02 12.00 


. . . 24.00 


A03 23.00 


. . . 46.00 


A04 25.50 


. . . 51.00 


A05 27.00 


. . . 54.00 


A06 29.50 


. . . 59.00 


A07 33.00 


. . . 66.00 


A08 36.00 


. . . 72.00 


A09 41.00 


. . . 82.00 


A10 44.00 


. . . 88.00 


All 47.00 


. . . 94.00 


A12 51.00 


. . 102.00 


A13 54.00 


. . 108.00 


A14 56.00 


. . 112.00 


A15 58.00 


. . 116.00 


A16 60.00 


. . 120.00 


A17 62.00 


. . 124.00 


A18 65.50 


. . 131.00 


A19 67.50 


. . 135.00 


A20 69.50 


. . 139.00 


A21 71.50 


. . 143.00 


A22 77.00 


. . 154.00 


A23 79.00 


. . 158.00 


A24 81.00 


. . 162.00 


A25 83.00 


. . 166.00 


A99 Contact NASA CASI 



U.S., Canada, 




Code & Mexico 


Foreigi 


E01 $101.00 


. . $202.00 


E02 109.50 


. . . 219.00 


E03 119.50 


. . . 238.00 


E04 128.50 


. . . 257.00 


E05 138.00 


. . . 276.00 


E06 146.50 


. . . 293.00 


E07 156.00 


. . . 312.00 


E08 165.50 


. . . 331.00 


E09 174.00 


. . . 348.00 


E10 183.50 


. . . 367.00 


Ell 193.00 


. . . 386.00 


E12 201.00 


. . . 402.00 


E13 210.50 


. . . 421.00 


E14 220.00 


. . . 440.00 


E15 229.50 


. . . 459.00 


E16 238.00 


. . . 476.00 


E17 247.50 


. . . 495.00 


E18 257.00 


. . . 514.00 


E19 265.50 


. . . 531.00 


E20 275.00 


. . . 550.00 


E21 284.50 


. . . 569.00 


E22 293.00 


. . . 586.00 


E23 302.50 


. . . 605.00 


E24 312.00 


. . . 624.00 


E99 Contact NASA CASI 



Payment Options 

All orders must be prepaid unless you are registered for invoicing or have a deposit account with the NASA CASI. 
Payment can be made by VISA, MasterCard, American Express, or Diner's Club credit card. Checks or money orders 
must be in U.S. currency and made payable to "NASA Center for AeroSpace Information." To register, please request 
a registration form through the NASA STI Help Desk at the numbers or addresses below. 

Handling fee per item is $1.50 domestic delivery to any location in the United States and $9.00 foreign delivery to 
Canada, Mexico, and other foreign locations. Video orders incur an additional $2.00 handling fee per title. 

The fee for shipping the safest and fastest way via Federal Express is in addition to the regular handling fee explained 
above — $5.00 domestic per item, $27.00 foreign for the first 1-3 items, $9.00 for each additional item. 

Return Policy 

The NASA Center for AeroSpace Information will replace or make full refund on items you have requested if we have 
made an error in your order, if the item is defective, or if it was received in damaged condition, and you contact CASI 
within 30 days of your original request. 



NASA Center for AeroSpace Information 
7121 Standard Drive 
Hanover, MD 21076-1320 

Rev. 7/98 



E-mail: help@sti.nasa.gov 
Fax: (301) 621-0134 
Phone: (301) 621-0390 



Federal Depository Library Program 

In order to provide the general public with greater access to U.S. Government publications, Congress 
established the Federal Depository Library Program under the Government Printing Office (GPO), 
with 53 regional depositories responsible for permanent retention of material, inter-library loan, and 
reference services. At least one copy of nearly every NASA and NASA-sponsored publication, 
either in printed or microfiche format, is received and retained by the 53 regional depositories. A list 
of the Federal Regional Depository Libraries, arranged alphabetically by state, appears at the very 
end of this section. These libraries are not sales outlets. A local library can contact a regional 
depository to help locate specific reports, or direct contact may be made by an individual. 

Public Collection of NASA Documents 

An extensive collection of NASA and NASA- sponsored publications is maintained by the British 
Library Lending Division, Boston Spa, Wetherby, Yorkshire, England for public access. The British 
Library Lending Division also has available many of the non-NASA publications cited in the STI 
Database. European requesters may purchase facsimile copy or microfiche of NASA and 
NASA- sponsored documents FIZ-Fachinformation Karlsruhe-Bibliographic Service, D-76344 
Eggenstein-Leopoldshafen, Germany and TIB— Technische Informationsbibliothek, P.O. Box 
60 80, D-30080 Hannover, Germany. 

Submitting Documents 

All users of this abstract service are urged to forward reports to be considered for announcement in 
the STI Database. This will aid NASA in its efforts to provide the fullest possible coverage of all 
scientific and technical publications that might support aeronautics and space research and 
development. If you have prepared relevant reports (other than those you will transmit to NASA, 
DOD, or DOE through the usual contract- or grant-reporting channels), please send them for 
consideration to: 

ATTN: Acquisitions Specialist 

NASA Center for AeroSpace Information 

7121 Standard Drive 

Hanover, MD 21076-1320. 

Reprints of journal articles, book chapters, and conference papers are also welcome. 

You may specify a particular source to be included in a report announcement if you wish; otherwise 
the report will be placed on a public sale at the NASA Center for AeroSpace Information. 
Copyrighted publications will be announced but not distributed or sold. 



Federal Regional Depository Libraries 



ALABAMA 

AUBURN UNIV. AT MONTGOMERY 
LIBRARY 

Documents Dept. 

7300 University Dr. 

Montgomery, AL 361 1 7-3596 

(205) 244-3650 Fax: (205) 244-0678 

UNIV. OF ALABAMA 

Amelia Gayle Gorgas Library 

Govt. Documents 

P.O. Box 870266 

Tuscaloosa, AL 35487-0266 

(205) 348-6046 Fax: (205) 348-0760 

ARIZONA 

DEPT. OF LIBRARY, ARCHIVES, 
AND PUBLIC RECORDS 

Research Division 

Third Floor, State Capitol 

1700 West Washington 

Phoenix, AZ 85007 

(602) 542-3701 Fax: (602) 542-4400 

ARKANSAS 

ARKANSAS STATE LIBRARY 

State Library Service Section 

Documents Service Section 

One Capitol Mall 

Little Rock, AR 72201-1014 

(501) 682-2053 Fax: (501) 682-1529 

CALIFORNIA 

CALIFORNIA STATE LIBRARY 

Govt. Publications Section 

P.O. Box 942837 - 914 Capitol Mall 

Sacramento, CA 94337-0091 

(916) 654-0069 Fax: (916) 654-0241 

COLORADO 

UNIV. OF COLORADO - BOULDER 

Libraries - Govt. Publications 

Campus Box 184 

Boulder, CO 80309-0184 

(303) 492-8834 Fax: (303) 492-1881 

DENVER PUBLIC LIBRARY 

Govt. Publications Dept. BSG 

1 357 Broadway 

Denver, CO 80203-2165 

(303) 640-8846 Fax: (303) 640-8817 

CONNECTICUT 

CONNECTICUT STATE LIBRARY 

231 Capitol Avenue 
Hartford, CT 06106 
(203) 566-4971 Fax: (203) 566-3322 

FLORIDA 

UNIV. OF FLORIDA LIBRARIES 

Documents Dept. 

240 Library West 

Gainesville, FL 32611-2048 

(904) 392-0366 Fax: (904) 392-7251 

GEORGIA 

UNIV. OF GEORGIA LIBRARIES 

Govt. Documents Dept. 

Jackson Street 

Athens, GA 30602-1 645 

(706) 542-8949 Fax: (706) 542-4144 

HAWAII 

UNIV. OF HAWAII 

Hamilton Library 

Govt. Documents Collection 

2550 The Mall 

Honolulu, HI 96822 

) 948-8230 Fax: (808) 956-5968 



IDAHO 

UNIV. OF IDAHO LIBRARY 

Documents Section 

Rayburn Street 

Moscow, ID 83844-2353 

(208) 885-6344 Fax: (208) 885-6817 

ILLINOIS 

ILLINOIS STATE LIBRARY 

Federal Documents Dept. 

300 South Second Street 

Springfield, IL 62701-1796 

(217) 782-7596 Fax: (217) 782-6437 



INDIANA 

INDIANA STATE LIBRARY 

Serials/Documents Section 
140 North Senate Avenue 
Indianapolis, IN 46204-2296 

(317) 232-3679 Fax: (317) 232-3728 

IOWA 

UNIV. OF IOWA LIBRARIES 

Govt. Publications 

Washington & Madison Streets 

Iowa City, IA 52242-1 166 

(319) 335-5926 Fax: (319) 335-5900 

KANSAS 

UNIV. OF KANSAS 

Govt. Documents & Maps Library 

6001 Malott Hall 

Lawrence, KS 66045-2800 

(913) 864-4660 Fax: (913) 864-3855 

KENTUCKY 

UNIV. OF KENTUCKY 

King Library South 

Govt. Publications/Maps Dept. 

Patterson Drive 

Lexington, KY 40506-0039 

(606) 257-3139 Fax: (606) 257-3139 

LOUISIANA 

LOUISIANA STATE UNIV. 

Middleton Library 

Govt. Documents Dept. 

Baton Rouge, LA 70803-3312 

(504) 388-2570 Fax: (504) 388-6992 

LOUISIANA TECHNICAL UNIV. 

Prescott Memorial Library 
Govt. Documents Dept. 
Ruston, LA 71 272-0046 

(318) 257-4962 Fax: (318) 257-2447 

MAINE 

UNIV. OF MAINE 

Raymond H. Fogler Library 

Govt. Documents Dept. 

Orono, ME 04469-5729 

(207) 581-1673 Fax: (207) 581-1653 

MARYLAND 

UNIV. OF MARYLAND - COLLEGE PARK 

McKeldin Library 
Govt. Documents/Maps Unit 
College Park, MD 20742 
(301)405-9165 Fax: (301)314-9416 

MASSACHUSETTS 

BOSTON PUBLIC LIBRARY 

Govt. Documents 
666 Boylston Street 
Boston, MA 021 17-0286 
(617)536-5400, ext. 226 
Fax: (617)536-7758 

MICHIGAN 

DETROIT PUBLIC LIBRARY 

5201 Woodward Avenue 

Detroit, Ml 48202-4093 

(313) 833-1025 Fax: (313) 833-0156 

LIBRARY OF MICHIGAN 

Govt. Documents Unit 

P.O. Box 30007 

717 West Allegan Street 

Lansing, Ml 48909 

(517) 373-1300 Fax: (517) 373-3381 

MINNESOTA 

UNIV. OF MINNESOTA 

Govt. Publications 

409 Wilson Library 

309 19th Avenue South 

Minneapolis, MN 55455 

(612) 624-5073 Fax: (612) 626-9353 

MISSISSIPPI 

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Typical Report Citation and Abstract 



O 19970001126 NASA Langley Research Center, Hampton, VA USA 

© Water Tunnel Flow Visualization Study Through Posts tall of 1.2 Novel Planform Shapes 

© Gatlin, Gregory M., NASA Langley Research Center, USA Neuhart, Dan H., Lockheed Engineering and Sciences Co., USA; 
© Mar. 1996; 130p; In English 
© Contract(s)/Grant(s): RTOP 505-68-70-04 

© Report No(s): NASA-TM-4663; NAS 1.15:4663; L-17418; No Copyright; Avail: CASI; A07, Hardcopy; A02, Microfiche 
© To determine the flow field characteristics of 12 planform geometries, a flow visualization investigation was conducted 

in the Langley 16- by 24-Inch Water Tunnel. Concepts studied included flat plate representations of diamond wings, twin 
bodies, double wings, cutout wing configurations, and serrated forebodies. The off-surface flow patterns were identified by 
injecting colored dyes from the model surface into the free-stream flow. These dyes generally were injected so that the local- 
ized vortical flow patterns were visualized. Photographs were obtained for angles of attack ranging from 10' to 50', and all 
investigations were conducted at a test section speed of 0.25 ft per sec. Results from the investigation indicate that the forma- 
tion of strong vortices on highly swept forebodies can improve poststall lift characteristics; however, the asymmetric bursting 
of these vortices could produce substantial control problems. A wing cutout was found to significantly alter the position of 
the forebody vortex on the wing by shifting the vortex inboard. Serrated forebodies were found to effectively generate multi- 
ple vortices over the configuration. Vortices from 65' swept forebody serrations tended to roll together, while vortices from 
40' swept serrations were more effective in generating additional lift caused by their more independent nature. 
© Author 

© Water Tunnel Tests; Flow Visualization; Flow Distribution; Free Flow; Planforms; Wing Profiles; Aerodynamic 
Configurations 



Key 



1. Document ID Number; Corporate Source 

2. Title 

3. Author(s) and Affiliation(s) 

4. Publication Date 

5. Contract/Grant Number(s) 

6. Report Number(s); Availability and Price Codes 

7. Abstract 

8. Abstract Author 

9. Subject Terms 



AERONAUTICAL 

ENGINEERING A Continuing Bibliography (Suppl. 387) 



NOVEMBER 13, 1998 



01 
AERONAUTICS 

19980227102 NASA Langley Research Center, Hampton, VA USA 

Aeronautical Engineering: A Continuing Bibliography, Supplment 385 

Oct. 16, 1998; 42p; In English 

Report No.(s): NASA/SP-1998-7037/SUPPL385; NAS 1.21:7037/SUPPL385; No Copyright; Avail: CASI; A03, Hardcopy; 

A01, Microfiche 

This supplemental issue of Aeronautical Engineering, A Continuing Bibliography with Indexes (NASA/SP-1998-7037) lists 
reports, articles, and other documents recently announced in the NASA STI Database. The coverage includes documents on the 
engineering and theoretical aspects of design, construction, evaluation, testing, operation, and performance of aircraft (including 
aircraft engines) and associated components, equipment, and systems. It also includes research and development in aerodynamics, 
aeronautics, and ground support equipment for aeronautical vehicles. Each entry in the publication consists of a standard biblio- 
graphic citation accompanied, in most cases, by an abstract. 
CASI 
Bibliographies; Aerodynamics; Aeronautical Engineering; Indexes (Documentation); Aircraft Design 

19980227425 Carnegie-Mellon Inst, of Research, Pittsburgh, PA USA 
Automated Inspection of Aircraft Final Report 

Alberts, C. J., Carnegie-Mellon Inst, of Research, USA; Carroll, C. W., Carnegie-Mellon Inst, of Research, USA; Kaufman, W. 
M., Carnegie-Mellon Inst, of Research, USA; Perlee, C. J., Carnegie-Mellon Inst, of Research, USA; Siegel, M. W., Carnegie- 
Mellon Inst, of Research, USA; Apr. 1998; 85p; In English 
Contract(s)/Grant(s): FAA94-G-018 

Report No.(s): AD-A350525; AAR-430; DOT/FAA/AR-97/69; No Copyright; Avail: CASI; A05, Hardcopy; A01, Microfiche 
This report summarizes the development of a robotic system designed to assist aircraft inspectors by remotely deploying non- 
destructive inspection (NDI) sensors and acquiring, processing, and storing inspection data. Carnegie Mellon University studied 
the task 9f aircraft inspection, compiled the functional requirements for an automated system to inspect skin fastener rows, and 
developed a conceptual design of an inspection robot. A prototype of the robotic inspection system (the Automated Nondestruc- 
tive Inspector or (ANDI) was developed. The first phase of system development resulted in a laboratory system that demonstrated 
the abilities to adhere to the surface of an aircraft panel and deploy a standard eddy-current sensor. The second phase of develop- 
ment included enhancing the mechanics, adding video cameras to the robot for navigation, and adding an on-board computer for 
low-level task sequencing. The second-phase system was subsequently demonstrated at the FAA's Aging Aircraft NDI Validation 
Center (AANC). During the final phase of development, emphasis was placed on the enhancement of the robot's navigational 
system through automated recognition of image features captured by the navigation cameras. A significant development effort 
remains to be accomplished before this robotic inspection technology is suitable for operational deployment. Outstanding devel- 
opment issues include: (1) reducing the weight of the robot so that it is more comfortable to lift and position on the aircraft; (2) 
improving the mechanical reliability and speed of the system; (3) minimizing the scratching of the skin surface by the suction cups 
and eddy-current sensors; (4) reduction or elimination of the umbilical cable; and (5) automation of the manually controlled opera- 
tions, to commercialize the technology, a new mechanical system would need to be designed and built incorporating the lessons 
of this work. 
DTIC 
Inspection; Aircraft Structures; Aircraft Maintenance; Robot Control; Robotics; Nondestructive Tests 



02 
AERODYNAMICS 

Includes aerodynamics of bodies, combinations, wings, rotors, and control surfaces; and internal flow in ducts and turbomachinery. 

19980221789 European Organization for the Safety of Air Navigation, Bretigny-sur-Orge, France 

User Manual for the Base of Aircraft Data (BAD A) 

Bos, A., European Organization for the Safety of Air Navigation, France; Mar. 1998; lOOp; In English 

Report No.(s): PB98-164312; EEC/NOTE-6/98-Rev-3.0; No Copyright; Avail: CASI; A05, Hardcopy; A02, Microfiche 

The BASE of Aircraft Data (BAD A) provides a set of ASCII files containing performance and operating procedure coeffi- 
cients for 151 different aircraft types. The coefficients include those used to calculate thrust, drag and fuel flow and those used 
to specify nominal cruise, climb and decent speed. User Manual for Revision 3.0 of BADA provides definitions of each of the 
coefficients and then explains the file formats. Instructions for remotely accessing the files via Internet are also given. 
NTIS 
Aerodynamic Drag; User Manuals (Computer Programs); Climbing Flight 

19980223077 NASA Ames Research Center, Moffett Field, CA USA 

Lift, Drag, Static Stability and Control Characteristics and Control Surface Panel Loads From Wind Tunnel Tests at 
Supersonic Speeds of Models of Two Versions of the B-70 Airplane 

Daugherty, James C, NASA Ames Research Center, USA; Green, Kendal H., NASA Ames Research Center, USA; Nelson, Rich- 
ard D., NASA Ames Research Center, USA; Oct. 1960; 148p; In English 
Report No.(s): NASA-TM-SX-396; X68-84368; No Copyright; Avail: CASI; A07, Hardcopy; A02, Microfiche 

Models of two versions of the North American B-70 airplane were tested at Mach numbers of 2.5 and 3.0 at a Reynolds num- 
ber of 5.0 million and at a Mach number of 3.5 at a Reynolds number of 4.5 million. Lift, drag, static-stability, and control charac- 
teristics were determined for both complete models, for various components, and for various component modifications. Flow 
visualization studies were made with both natural and fixed boundary-layer transition. The data for the best configuration tested 
were extrapolated to give an all-turbulent lift-drag ratio for a Reynolds number of 110 million at a Mach number of 3.0. 
Author 

Wind Tunnel Tests; B-70 Aircraft; Boundary Layer Transition; Flow Visualization; Mach Number; Reynolds Number; Supersonic 
Speed; Lift; Drag; Static Stability; Controllability; Dynamic Characteristics 

19980223577 NASA Langley Research Center, Hampton, VA USA 

Analytic Study of Induced Pressure on Long Bodies of Revolution with Varying Nose Bluotoess at Hypersonic Speeds 

VanHise, Vernon, NASA Langley Research Center, USA; 1961; 20p; In English 

Report No.(s): NASA-TR-R-78; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Pressure distributions and shock shapes for a series of cylindrical afterbodies having nose fineness ratios from 0.4 to 4 have 
been calculated by using the method of characteristics for a perfect gas. The fluid mediums investigated were air and helium and 
the Mach number range was from 5 to 40. Flow parameters obtained from blast-wave analogy gave good correlations of blunt-nose 
induced pressures and shock shapes. Experimental results are found to be in good agreement with the characteristic calculations. 
The concept of hypersonic similitude enables good correlation of the results with respect to body shape, Mach number, and ratio 
of specific heats. 
Author 

Hypersonic Speed; Pressure Distribution; Cylindrical, Bodies; Bodies of Revolution; Afterbodies; Hypersonics; Method of Char- 
acteristics; Blunt Bodies; Shock Waves 

19980223578 NASA Ames Research Center, Moffett Field, CA USA 

Effects of Outboard Thickened and Blunted Leading Edges on the Wave Drag of a 45 Degree Swept- Wing and Body Com- 
bination 

Holdaway, George H., NASA Ames Research Center, USA; Lazzeroni, Frank A., NASA Ames Research Center, USA; Hatfield, 
Elaine W., NASA Ames Research Center, USA; Aug. 1959; 34p; In English 
Report No.(s): NASA-TM-X-27; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An investigation to evaluate the effects of thickened and blunted leading-edge modifications on the wave drag of a swept wing 
has been made at Mach numbers from 0.65 to 2.20 and at a Reynolds number of 2,580,000 based on the mean aerodynamic chord 
of the basic wing. Two leading-edge designs were investigated and they are referred to as the thickened and the blunted modifica- 
tions although both sections had equally large leading-edge radii. The thickened leading edge was formed by increasing the thick- 



ness over the forward 40 percent of the basic wing section. The blunted modification was formed by reducing the wing chords 
about 1 percent and by increasing the section thickness slightly over the forward 6 percent of the basic section in a manner to keep 
the wing sweep and volume essentially equal to the respective values for the basic wing. The basic wing had an aspect ratio of 
3, a leading-edge sweep of 45 deg., a taper ratio of 0.4, and NACA 64AO06 sections perpendicular to a line swept back 39.45 
deg., the quarter-chord line of these sections. Test results indicated that the thickened modification resulted in an increase in zero- 
lift drag coefficient of from 0.0040 to 0.0060 over values for the basic model at Mach numbers at which the wing leading edge 
was sonic or supersonic. Although drag coefficients of both the basic and thickened models were reduced at all test Mach numbers 
by body indentations designed for the range of Mach numbers from 1 .00 to 2.00, the greater drag of the thickened model relative 
to that of the basic model was not reduced. The blunted model, however, had less than one quarter of the drag penalty of the thick- 
ened model relative to the basic model at supersonic leading -edge conditions (M greater or equal to root-2). 
Author 

Swept Wings; Blunt Bodies; Aerodynamic Coefficients; Airfoil Profiles; Leading Edges; Thickness; Wave Drag; Wind Tunnel 
Tests; Wind Tunnel Models 



19980223579 NASA Langley Research Center, Hampton, VA USA 

Large Angle Motion Tests, Including Spins, of a Free-Flying Radio-Controlled 0.13-Scale Model of a Twin-Jet Swept- 

Wiiig Fighter Airplane 

Burk, Sanger M., Jr., NASA Langley Research Center, USA; Libbey, Charles E., NASA Langley Research Center, USA; 1961; 

42p; In English 

Report No.(s): NASA-TM-SX-445; L-1192; N-AM-50; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An investigation has been conducted with a free-flying radio-controlled 0. 13-scale model of a twin-jet swept-wing fighter 
airplane to determine the tendency of this design to enter spins and to evaluate the nature of the spin obtained from post-stall 
motions. The test results indicate that it may be difficult to obtain a developed spin on the airplane, particularly the flat-type spin. 
Two types of erect developed spins will be possible; one will be flat and fast rotating from which recovery may not be obtained 
and the other will be steeper and oscillatory from which recoveries will be satisfactory. Controls will be effective for satisfactory 
termination of the post-stall gyrations obtained. The recommended recovery technique from both post-stall gyrations and devel- 
oped spins will be movement of the rudder to oppose the yawing rotation and simultaneous movement of the ailerons to with the 
rotation (stick right when turning to the right). When recovery is imminent, the stick should be moved longitudinally to neutral. 
It is recommended that the spin not be allowed to develop fully on this airplane. The developed-spin results obtained in the inves- 
tigation were in good agreement with spin-tunnel results. 
Author 

Fighter Aircraft; Aircraft Spin; Spin Dynamics; Aerodynamic Stalling; Wind Tunnel Tests; Wind. Tunnel Models; Scale Models; 
Swept Wings; Control Stability; Maneuvers; Aircraft Stability 



19980223580 NASA Langley Research Center, Hampton, VA USA 

Free-Spinning-Tunnel Investigation of a 1/30 Scale Model of a Twin-Jet-Swept- Wing Fig 

Bowman, James S., Jr., NASA Langley Research Center, USA; Healy, Frederick M., NASA Langley Research Center, USA; 

1960; 22p; In English 

Report No.(s): NASA-TM-SX-446; L-1191; N5154; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An investigation has been made in the Langley 20-foot free-spinning tunnel to determine the erect and inverted spin and 
recovery characteristics of a 1/30-scale dynamic model of a twin-jet swept-wing fighter airplane. The model results indicate that 
the optimum erect spin recovery technique determined (simultaneous rudder reversal to full against the spin and aileron deflection 
to full with the spin) will provide satisfactory recovery from steep-type spins obtained on the airplane. It is considered that the 
air-plane will not readily enter flat-type spins, also indicated as possible by the model tests, but developed-spin conditions should 
be avoided in as much as the optimum recovery procedure may not provide satisfactory recovery if the airplane encounters a flat- 
type developed spin. Satisfactory recovery from inverted spins will be obtained on the airplane by neutralization of all controls. 
A 30-foot- diameter (laid-out-flat) stable tail parachute having a drag coefficient of 0.67 and a towline length of 27.5 feet will be 
satisfactory for emergency spin recovery. 
Author 

Aircraft Spin; Control Stability; Spin Dynamics; Wind Tunnel Tests; Wind Tunnel Models; Aerodynamic Stalling; Fighter Aircraft; 
Scale Models; Swept Wings; Aircraft Stability 



19980223582 NASA Langley Research Center, Hampton, VA USA 

Summary of Results Obtained in Full-Scale Tiinnel Investigation of the Ryan Flex- Wing Airplane 

Johnson, Joseph L., Jr., NASA Langley Research Center, USA; Hassell, James L., Jr., NASA Langley Research Center, USA; Aug. 

14, 1962; 40p; In English 

Report No.(s): NASA-TM-SX-727; L-3093; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The performance and static stability and control characteristics of the Ryan Flex-Wing airplane were determined in an inves- 
tigation conducted in the Langley full-scale tunnel through an angle-of-attack range of the keel from about 14 to 44 deg. for pow- 
er-on and -off conditions. Comparisons of the wind-tunnel data with flight-test data obtained with the same airplane by the Ryan 
Aeronautical Company were made in a number of cases. 
Author 
Wind Tunnel Tests; Static Stability; Aircraft Performance; Wings; Longitudinal Stability; Aircraft Stability; Flight Characteristics 

19981)223585 NASA Langley Research Center, Hampton, VA USA 

An Investigation of the Influence of Body Size and Indentation Asymmetry of the Effectiveness of Body Indentation in 

Combination with a Cambered Wing 

Patterson, James C, Jr., NASA Langley Research Center, USA; Loving, Donald L., NASA Langley Research Center, USA; Feb. 

1961; 38p; In English 

Report No.(s): NASA-TM-X-427; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An investigation has been made of a 450 sweptback cambered wing in combination with an unindented body and a body sym- 
metrically indented with respect to its axes designed for a Mach number of 1.2. The ratio of body frontal area to wing planform 
area was 0.08 for these wing-body combinations. In order to determine the influence of body size on the effectiveness of indenta- 
tion, the test data have been compared with previously obtained data for similar configurations having a ratio of body frontal area 
to wing planform area of 0.04. Also, in order to investigate the relative effectiveness of indentation asymmetry, a specially 
indented body designed to account for the wing camber and also designed for a Mach number of 1.2 has been included in these 
tests. The investigation was conducted in the Langley 8-Foot Tunnels Branch at Mach numbers from 0.80 to 1.43 and a Reynolds 
number of approximately 1.85 x 10(exp 6), based on a mean aerodynamic chord length of 5.955 inches. The data indicate that 
the configurations with larger ratio of body frontal area to wing planform area had smaller reductions in zero-lift wave drag associ- 
ated with body indentation than the configurations with smaller ratio of body frontal area to wing planform area. The 0.08-area-ra- 
tio configurations also had correspondingly smaller increases in the values of maximum lift-drag ratio than the 0.04-area-ratio 
configurations. The consideration of wing camber in the body indentation design resulted in a 35.5-percent reduction in zero-lift 
wave drag, compared with a 21.5-percent reduction associated with the symmetrical indentation, but had a negligible effect on 
the values of maximum lift-drag ratio. 
Author 
Body-Wing Configurations; Lift Drag Ratio; Indentation; Cambered Wings; Airfoil Profiles; Sweptback Wings 

19980223586 NASA Dryden Flight Research Center, Edwards, CA USA 

Preliminary Fuii-Scale Power-Off Drag of the X-1S Airplane for Mach Numbers from 0.7 to 3.1 
Saltzman, Edwin J., NASA Dryden Flight Research Center, USA; Dec. 1960; 26p; In English 
Report No.(s): NASA-TM-X-430; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Drag characteristics have been obtained for the X-15 airplane during unpowered flight. These data represent a Mach number 
range from about 0.7 to 3.1 and a Reynolds number range from 13.9 x 10(exp 6) to 28 x 10(exp 8), based on the mean aerodynamic 
chord. The full-scale data are compared with estimates compiled from several wind-tunnel facilities. The agreement between 
wind-tunnel and full-scale supersonic drag, uncorrected for Reynolds number effects, is reasonably close except at low supersonic 
Mach numbers where the flight values are significantly higher. 
Author 
X-15 Aircraft; Aerodynamic Drag; Mach Number; Airfoil Profiles; Supersonic Drag; Wind Tunnel Tests 

19980223590 NASA Langley Research Center, Hampton, VA USA 

Several Methods for Aerodynamic Reduction of Static-Pressure Sensing Errors for Aircraft at Subsonic, Near-Sonic, and 

Low Supersonic Speeds 

Ritchie, Virgil S., NASA Langley Research Center, USA; 1959; 28p; In English 

Report No.(s): NASA-TR-R-18; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Tests were conducted in transonic wind tunnels to investigate and verify experimentally methods for aerodynamically reduc- 
ing errors due to sensor position, bow-wave passage, and angle of attack. The results indicated that aerodynamic devices of simple 



design may be employed to reduce errors in sensing static pressures to less than 0.5 percent at Mach numbers from about 0.40 

to 1.15. 

Author 

Pressure Reduction; Aerodynamics; Static Pressure; Angle of Attack; Errors; Detection; Bow Waves 

19989223594 NASA Langley Research Center, Hampton, VA USA 

Ordinate's and Theoretical Pressure-Distribution Data for NACA 6~ and 6A -Series Airfoil Sections with Thicknesses from 

2 to 21 and From 2 to 15 Percent Chord, Respectively 

Patterson, Elizabeth W., NASA Langley Research Center, USA; Braslow, Albert L., NASA Langley Research Center, USA; 1961; 

88p; In English 

Report No.(s): NASA-TR-R-84; No Copyright; Avail: CASI; A05, Hardcopy; A01, Microfiche 

Information is presented with which ordinates can be easily obtained for any thickness from 2 to 21 percent chord for NACA 
63-, 64-, and 65-series airfoil sections and from 2 to 15 percent chord for NACA 63A-, 64- A, series airfoil sections. In addition, 
data required for estimation of the theoretical pressure distributions of any of these airfoils are included. 
Author 
Airfoils; Pressure Distribution; Airfoil Profiles; Aircraft Design 

19980223692 NASA Ames Research Center, Moffett Field, CA USA 

Lift, Drag, and Pitching Moments of an Arrow Wing Having 81) Degree of Sweepback at Mach Numbers from 2,48 to 3.51 

and Reynolds Numbers up to 11.0 Million 

Hopkins, Edward J., NASA Ames Research Center, USA; Jillie, Don W., NASA Ames Research Center, USA; Levin, Alan D., 

NASA Ames Research Center, USA; Aug. 1959; 46p; In English 

Report No.(s): NASA-TM-X-22; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Measurements were made of the lift, drag, and pitching moments on an arrow wing (taper ratio of zero) having an aspect ratio 
of 1.4 and a leading-edge sweepback of 80 (degrees). The wing was designed to have a subsonic leading-edge and a Clark-Y airfoil 
with a thickness ratio of 12 percent of the chord perpendicular to the wing leading edge. The wing was tested both with and without 
the wing tips bent upward in an attempt to alleviate possible flow separation in the vicinity of the wing tips. Small jets of air were 
used to fix transition near the wing leading edge. Force results are presented for Mach numbers of 2.48, 2.75, 3.04, 3.28, and 3.51 
at Reynolds numbers of 3.5 and 9.0 million and for a Mach number of 3.04 at a Reynolds number of 11.0 million. The measured 
aerodynamic characteristics are compared with those estimated by linear theory. The maximum lift-drag ratio measured was much 
less than that predicted. This difference is attributed to lack of full leading-edge thrust and to the experimental lift-curve slope 
being about 20 percent below the theoretical value. 
Author 

Lift; Aerodynamic Drag; Pitching Moments; Separated Flow; Wind Tunnel Tests; Sweptbaclc Wings; Boundary Layer Separation; 
Arrow Wings 

19980223693 NASA Langley Research Center, Hampton, VA USA 

Effect of Multiple-Jets Exits on the Base Pressure of a Simple Wing-Body Combination at Mach Numbers of 0.6 to 1.27 
Cubbage, James M., Jr., NASA Langley Research Center, USA; Aug. 1959; 38p; In English 
Report No.(s): NASA-TM-X-25; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An investigation has been conducted at Mach numbers of 0.6 to 1 .27 to determine the effect of multiple-jet exits on the base 
pressure of a simple wing-body combination. The design Mach number of the nozzles ranged from 1 to 3 at jet exit diameters equal 
to 36.4 to 75 percent of the model thickness. Jet total-pressure to free-stream static -pressure ratios ranged from 1 (no flow) to 34.2. 
The results show that the variation of base pressure coefficient with jet pressure ratio for the model tested was similar to that 
obtained for single nozzles in bodies of revolution in other investigations. As in the case for single jets the base pressure coefficient 
for the present model became less negative as the jet exit diameter increased. For a constant throat diameter and an assumed sched- 
ule of jet pressure ratio over the speed range of these tests, nozzle Mach number had only a small effect on base pressure coefficient. 
Author 
Body-Wing Configurations; Base Pressure; Transonic Speed; Jet Exhaust; Free Flow; Exhaust Nozzles 

19980223695 NASA Langley Research Center, Hampton, VA USA 

A Supersonic Area Rule and an Application to the Design of a Wing-Body Combination with High Lift-Drag Ratios 
Whitcomb, Richard T., NASA Langley Research Center, USA; Sevier, John R., Jr., NASA Langley Research Center, USA; 1960; 
16p; In English 



Report No.(s): NASA-TR-R-72; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A concept for interrelating the wave drags of wing-body combinations at supersonic speeds with axial developments of cross- 
sectional area is presented. A swept-wing-indented-body combination designed on the basis of this concept to have significantly 
improved maximum lift-drag ratios over a range of transonic and moderate supersonic speeds is described. Experimental results 
have been obtained for this configuration at Mach numbers from 0.80 to 2.01. Maximum lift-drag ratios of approximately 14 and 
9 were measured at Mach numbers of 1.15 and 1.41, respectively. 
Author 
Body-Wing Configurations; Supersonic Speed; Wave Drag; Transonic Speed; Lift Drag Ratio 

19980223699 NASA Langley Research Center, Hampton, VA USA 

Calculated Effects of Body Shape oh the Bow-Shock Overpressures io the Far Field of Bodies in Supersonic Flow 

Lansing, Donald L., NASA Langley Research Center, USA; 1960; 16p; In English 

Report No.(s): NASA-TR-R-76; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A theory for the supersonic flow about bodies in uniform flight in a homogeneous medium is reviewed and an integral which 
expresses the effect of body shape upon the flow parameters in the far field is reduced to a form which may be readily evaluated 
for arbitrary body shapes. This expression is then used to investigate the effect of nose angle, fineness ratio, and location of maxi- 
mum body cross section upon the far-field pressure jump across the bow-shock of slender bodies. Curves are presented showing 
the variation of the shock strength with each of these parameters. It is found that, for a wide variety of shapes having equal fineness 
ratios, the integral has nearly a constant value. 
Author (revised) 

Slender Bodies; Bow Waves; Sonic Booms; Supersonic Flight; Jet Aircraft Noise; Aircraft Structures; Aerodynamic Noise; Air- 
craft Design; Overpressure 

19980223614 NASA Langley Research Center, Hampton, VA USA 

Longitudinal Aerodynamic Characteristics of a Wing-Body -Tail Model Having a Highly Tapered, Cambered 45 degree 

Swept Wing of Aspect Ratio 4 at Transonic Speeds 

West, F. E., Jr., NASA Langley Research Center, USA; Nov. 1959; 32p; In English 

Report No.(s): NASA-TM-X-130; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The longitudinal aerodynamic characteristics of a wing-body -horizontal-tail configuration designed for efficient perfor- 
mance at transonic speeds has been investigated at Mach numbers from 0.80 to 1.03 in the Langley 16-foot transonic tunnel. The 
effect of adding an outboard leading-edge chord-extension to the highly tapered 45 deg. swept wing was also obtained. The aver- 
age Reynolds number for this investigation was 6.7 x 10(exp 6) based on the wing mean aerodynamic chord. The relatively low 
tail placement as well as the addition of a chord-extension achieved some alleviation of the pitchup tendencies of the wing-fuselage 
configuration. The maximum trimmed lift-drag ratio was 16.5 up to a Mach number of 0.9, with the moment center located at the 
quarter-chord point of the mean aerodynamic chord. For the untrimmed case, the maximum lift-drag ratio was approximately 19.5 
up to a Mach number of 0.9. 
Author 

Body-Wing and Tail Configurations; Aircraft Design; Swept Wings; Wind Tunnel, Tests; Transonic Speed; Cambered Wings; Aero- 
dynamic Characteristics; Aerodynamic Balance 

19980223615 NASA Langley Research Center, Hampton, VA USA 

Aerodynamic Characteristics at Mach Number 2,05 of a Series of Highly Swept Arrow Wings Employing Various Degrees 

of Twist and Camber 

Carlson, Harry W., NASA Langley Research Center, USA; Oct. 1960; 36p; In English 

Report No.(s): NASA-TM-X-332-1; L-876; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A series of arrow wings employing various degrees of twist and camber were tested in the Langley 4- by 4-foot supersonic 
pressure tunnel. Aerodynamic forces and moments in pitch were measured at a Mach number of 2.05 and at a Reynolds number 
of 4.4 x 10(exp 6) based on the mean aerodynamic chord. Three of the wings, having a leading-edge sweep angle of 70 deg. and 
an aspect ratio of 2.24, were designed to produce a minimum drag (in comparison with that produced for other wings in the family) 
at lift coefficients of 0. 0.08, and 0.16. A fourth and a fifth wing, having a 75 deg. swept leading edge and an aspect ratio of 1.65, 
were designed for lift coefficients of and 0. 16, respectively. A 70 deg. swept arrow wing with twist and camber designed for 
an optimum loading at a lift coefficient considerably less than that for maximum lift-drag ratio gave the highest lift-drag ratio of 
all the wings tested a value of 8.8 compared with a value of 8.1 for the corresponding wing without twist and camber. Two twisted 
and cambered wings designed for optimum loading at the lift coefficient for maximum lift-drag ratio gave only small increases 



in maximum lift-drag ratios over that obtained for the corresponding flat wings. However, in all cases, the lift-drag ratios obtained 

were far below the theoretical estimates. 

Author 

Aerodynamic Characteristics; Supersonic Speed; Arrow Wings; Swept Wings; Twisted Wings; Cambered Wings; Wind Tunnel 
Tests; Aerodynamic Coefficients; Aerodynamic Configurations 

19980223945 NASA Langley Research Center, Hampton, VA USA 

The Drag Coefficient of Parabolic Bodies of Revolution Operating at Zero Cavitation Number and Zero Angle of Yaw 

Johnson, Virgil E., Jr., NASA Langley Research Center, USA; Rasnick, Thomas A., NASA Langley Research Center, USA; 1961; 

20p; In English 

Report No.(s): NASA-TR-R-86; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The form-drag coefficient of parabolic bodies of revolution with fineness ratios greater than 1 operating at zero angle of yaw 
and zero cavitation number is determined both theoretically and experimentally. Agreement between theory and experiment is 
very good, The theoretical form-drag coefficient of paraboloids is about half the form-drag coefficient of cones of comparable 
fineness ratio. 
Author 
Aerodynamic Drag; Bodies of Revolution; Parabolic Bodies; Aerodynamic Coefficients; Cavitation Flow 

19981)223966 NASA Langley Research Center, Hampton, VA USA 

Characteristics of a Model of a Proposed Six-Engine Hull-Type Seaplane Designed for Super- 



Wornom, Dewey E., NASA Langley Research Center, USA; Mar. 1960; 40p; In English 
Report No.(s): NASA-TM-X-246; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Force tests of a model of a proposed six -engine hull-type seaplane were performed in the Langley 8-foot transonic pressure 
tunnel. The results of these tests have indicated that the model had a subsonic zero-lift drag coefficient of 0.0240 with the highest 
zero-lift drag coefficient slightly greater than twice the subsonic drag level. Pitchup tendencies were noted for subsonic Mach 
numbers at relatively high lift coefficients. Wing leading-edge droop increased the maximum lift-drag ratio approximately 8 per- 
cent at a Mach number of 0.80 but this effect was negligible at a Mach number of 0.90 and above. The configuration exhibited 
stable lateral characteristics over the test Mach number range. 
Author 

Transonic Speed; Aerodynamic Coefficients; Aerodynamic Characteristics; Seaplanes; Supersonic Flight; Wind Tunnel Tests; 
Wind Tunnel Models; Aerodynamic Configurations 

19980223978 NASA Langley Research Center, Hampton, VA USA 

Full-Scale Wind-Tunnel Investigation of the Drag Characteristics of an HU2K Helicopter Fuselage 

Scallion, William I., NASA Langley Research Center, USA; 1963; 34p; In English 

Report No.(s): NASA-TM-SX-848; L-3338; N-AM-110; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An investigation was conducted in the Langley full-scale tunnel to determine the drag characteristics of the HU2K helicopter 
fuselage. The effects of body shape, engine operation, appendages, and leakage on the model drag were determined. The results 
of the tests showed that the largest single contribution to the parasite drag was that of the rotor hub installation which produced 
about 80 percent of the drag of the sealed and faired production body. Fairings on the rotor hub and blade retentions, or a cleaned-up 
hub and retentions, appeared to be the most effective single modifications tested. The total drag of all protuberances and air leakage 
also contributed a major part of the drag - an 83-percent increase over the drag of the sealed and faired production body. An addi- 
tional increment of drag was caused by the basic shape of the fuselage - 19 percent more than the drag obtained when the fuselage 
shape was extensively refaired. Another sizable increment of drag was caused by the engine oil-cooler exit which gave a drag of 
8 percent of that of the sealed and faired production body. 
Author 
Wind Tunnel Tests; Helicopters; Fuselages; Rotary Wings; Fairings; Hubs; Protuberances; Aerodynamic Drag 

19980223979 NASA Langley Research Center, Hampton, VA USA 

Aerodynamic Characteristics of the Pershing Missile During Separation of its Three Stages 

McShera, John T., NASA Langley Research Center, USA; Townsend, Quwatha S., NASA Langley Research Center, USA; 1961; 

94p; In English 

Report No.(s): NASA-TM-SX-524; L-1360; A-AM-45; No Copyright; Avail: CASI; A05, Hardcopy; A01, Microfiche 



An investigation to determine the aerodynamic characteristics of the first stage of a Pershing missile during separation from 
the second stage and of the second stage during separation from the reentry body has been made. The tests were conducted over 
a Mach number range from 1.70 to 4.65 and at Reynolds numbers of 3 x 10(exp 6) and 6 x 10(exp 6). The horizontal separation 
distance between the two stages was varied from to 4 body diameters and the vertical separation distance was varied from to 
0.5 body diameters. The angle of attack of the forward stage was varied from deg. to 6 deg., the angle of attack of stage one 
relative to the second stage was varied from deg. to 5 deg., and the angle of attack of the second stage relative to the reentry body 
was varied from deg. to 3 deg. Forces and moments were measured on the separated stage and the base pressure was measured 
on the forward stage. Tare forces caused by the presence of the support strut attached to the forward stage were approximated by 
making tests of the model inverted with and without an image strut. The results of this investigation are presented in coefficient 
form in tables without analysis. 
Author 
Aerodynamic Characteristics; Pershing Missile; Supersonic Speed; Missile Tests; Missile Trajectories 

19980223980 NASA Ames Research Center, Moffett Field, CA USA 

Investigation at Mach Numbers of 0.60 to 3.50 of Blended Wing-Body Combinations with Cambered and Twisted Wings 

with Diamond, Delta and Arrow Plan Forms 

Holdaway, George H., NASA Ames Research Center, USA; Mellenthin, Jack A., NASA Ames Research Center, USA; Oct. 1960; 

80p; In English 

Report No.(s): NASA-TM-X-390; No Copyright; Avail: CASI; A05, Hardcopy; A01, Microfiche 

This investigation is a continuation of the experimental and theoretical evaluation of blended wing-body combinations. The 
basic diamond, delta, and arrow plan forms which had an aspect ratio of 2 with leading-edge sweeps of 45.00 deg., 59.04 deg., 
and 70.82 deg. and trailing edge of -45.00 deg., -18.43 deg., and 41.19 deg., respectively, are used herein as standards for evaluat- 
ing the effects of camber and warp. The wing thickness distributions were computed by varying the section shape along with the 
body radii (blending process) to match the prescribed area distribution and wing plan form. The wing camber and warp were com- 
puted to try to obtain nearly elliptical spanwise and chordwise load distributions for each plan form and thus to obtain low drag 
due to lift for a range of Mach numbers for which the velocities normal to the wing leading edge are subsonic. Elliptical chordwise 
load distributions were not possible for the plan forms and design conditions selected, so these distributions were somewhat differ- 
ent for each plan form. The models were tested with transition fixed at Mach numbers from 0.60 to 3.50 and at Reynolds numbers, 
based on the mean aerodynamic chord of the wing, of roughly 4,000,000 to 9,000,000. At speeds where the velocities normal to 
the wing leading edges were supersonic, an increase in the experimental wave-drag coefficients due to camber and twist was evi- 
dent, but this penalty decreased with increased sweep. Thus the minimum wave-drag coefficients for the cambered arrow model 
were almost identical with the zero-lift wave- drag coefficients for the uncambered arrow model at all test Mach numbers. 
Author 

Body-Wing Configurations; Cambered Wings; Twisted Wings; Aerodynamic Coefficients; Airfoil Profiles; Subsonic Speed; Tran- 
sonic Speed; Supersonic Speed 

19981)223994 NASA Ames Research Center, Moffett Field, CA USA 

Large-Scale Wind-Tiiiine! Tests of a Wingless Vertical Take-Off and Landing Aircraft: Preliminary Results 

Koenig, David G., NASA Ames Research Center, USA; Brady, James A., NASA Ames Research Center, USA; Oct. 1960; 44p; 

In English 

Report No.(s): NASA-TN-D-326; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Large-scale wind-tunnel tests were made of a wingless vertical take-off and landing aircraft at zero sideslip to determine per- 
formance and longitudinal stability and control characteristics at airspeeds from to 70 knots. Roll control and rudder effective- 
ness were also obtained. Limitations in the propulsion system restricted the lift for which level flight could be simulated to 
approximately 1500 pounds. Test variables with roll control and rudder undeflected were airspeed, vane setting, angle of attack, 
elevator deflection, and power. In most of the tests angle of attack, elevator, and power were varied individually while the other 
four parameters were held constant at previously determined values required for simulating trimmed level flight. The majority 
of the tests were made with power on and tail on at airspeeds between 20 and 70 knots. However, a limited number of data were 
obtained for the following conditions: (1) at zero velocity, horizontal tail on, power on; (2) at forward velocity, tail off and power 
on; and (3) at forward velocity, tail on, but with power off. 
Author 

Wind Tunnel Tests; Longitudinal Stability; Aerodynamic Characteristics; Aerodynamic Balance; Lateral Control; Vertical Land- 
ing; Aerodynamic Configurations; Wind Tunnel Models; Scale Models 



19980223995 NASA Langley Research Center, Hampton, VA USA 

Aerodynamic Characteristics at Mach Numbers from to 1.6 to 2.8 of 74 deg. Swept Arrow Wings with and without Camber 
and Twist 

Hasson, Dennis R, NASA Langley Research Center, USA; Fichter, Ann B., NASA Langley Research Center, USA; Wong, Nor- 
man, NASA Langley Research Center, USA; Sep. 1959; 38p; In English 
Report No.(s): NASA-TM-X-8; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An investigation has been conducted to determine the lift, drag, and pitching-moment characteristics of a cambered and 
twisted arrow wing and an uncambered and untwisted arrow wing. The cambered and twisted wing was designed to give a high 
value of maximum lift-drag ratio at a lift coefficient of 0. 1 and at a Mach number of 2.50. Each wing had a leading-edge sweep 
of 74 deg., an aspect ratio of 1.6, a taper ratio of 0, and a notch ratio of 0.714. A 3-percent-streamwise biconvex thickness distribu- 
tion was centered on the mean camber surface of both wings. Tests were conducted at Mach numbers from 1.6 to 2.8 through a 
range of angle of attack from -6 deg. to 14 deg. The Reynolds number based on mean aerodynamic chord was 5.0 x 10( exp 6) 
for all tests. The maximum lift-drag ratio at the design Mach number for the cambered and twisted wing was 7.85 and, thus, was 
below the theoretically predicted value of 9. 10. In addition, the cambered and twisted wing had only slightly higher values of maxi- 
mum lift-drag ratio throughout the test Mach number range than the uncambered and untwisted wing. With the moment reference 
centers at 0.565 wing mean aerodynamic chord, both wings were slightly unstable longitudinally at low lift coefficients. For lift 
coefficients greater than about 0.1, the instability became more marked. These characteristics were obtained at all Mach numbers 
at which tests were made. 
Author 

Aerodynamic Characteristics; Aerodynamic Coefficients; Arrow Wings; Cambered Wings; Supersonic Speed; Swept Wings; 
Twisted Wings 

19980227077 NASA Ames Research Center, Moffett Field, CA USA 

Large-Scale Wind-Tunnel Tests of an Airplane Model with an Unswept, Tilt Wing of Aspect Ratio 5.5, and with Four Pro- 
pellers and Blowing Flaps 

Weiberg, James A., NASA Ames Research Center, USA; Holzhauser, Curt A., NASA Ames Research Center, USA; Jun. 1961; 
34p; In English 
Report No.(s): NASA-TN-D-1034; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Tests were made of a large-scale tilt-wing deflected-slipstream VTOL airplane with blowing-type BLC trailing-edge flaps. 
The model was tested with flap deflections of deg. without BLC, 50 deg. with and without BLC, and 80 deg. with BLC for wing- 
tilt angles of 0, 30, and 50 deg. Included are results of tests of the model equipped with a leading-edge flap and the results of tests 
of the model in the presence of a ground plane. 
Author 

Wind Tunnel Tests; Unswept Wings; Externally Blown Flaps; Boundary Layer Control; Aircraft Models; Trailing Edge Flaps; 
Vertical Takeoff Aircraft; Aerodynamic Coefficients; Tilt Wing Aircraft 

19980227081 NASA Langley Research Center, Hampton, VA USA 

Transonic Aerodynamic Loading Characteristics of a Wing-Body-Tail Combination Having a 52.5 deg. Sweptback Wing 
of Aspect Ratio 3 With Conical Wing Camber and Body Indentation for a Design Mach Number of Square Moot of 2 
Cassetti, Marlowe D., NASA Langley Research Center, USA; Re, Richard J., NASA Langley Research Center, USA; Igoe, Wil- 
liam B., NASA Langley Research Center, USA; Oct. 1961; 102p; In English 
Report No.(s): NASA-TN-D-971; No Copyright; Avail: CASI; A06, Hardcopy; A02, Microfiche 

An investigation has been made of the effects of conical wing camber and body indentation according to the supersonic area 
rule on the aerodynamic wing loading characteristics of a wing-body-tail configuration at transonic speeds. The wing aspect ratio 
was 3, taper ratio was 0.1, and quarter-chord-line sweepback was 52.5 deg. with 3 -percent -thick airfoil sections. The tests were 
conducted in the Langley 16-foot transonic tunnel at Mach numbers from 0.80 to 1.05 and at angles of attack from deg. to 14 
deg., with Reynolds numbers based on mean aerodynamic chord varying from 7 x 10(exp 6) to 8 x 10(exp 6). Conical camber 
delayed wing-tip stall and reduced the severity of the accompanying longitudinal instability but did not appreciably affect the 
spanwise load distribution at angles of attack below tip stall. Body indentation reduced the transonic chordwise center-of -pressure 
travel from about 8 percent to 5 percent of the mean aerodynamic chord. 
Author 

Aerodynamic Loads; Body-Wing and Tail Configurations; Wind Tunnel Tests; Wing Loading; Sweptback Wings; Supersonic 
Speed; Conical Camber ; Airfoil Profiles; Aerodynamic Characteristics 



19980227085 NASA Langley Research Center, Hampton, VA USA 
teristics at Several Stations on a I 



Church, James D., NASA Langley Research Center, USA; Cremin, Joseph W., NASA Langley Research Center, USA; Nov. 1959; 

26p; In English 

Report No.(s): NASA-TM-X-59; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An investigation of the upwash characteristics at several longitudinal stations along and above the surface of a blunted cone- 
frustum-cylinder model has been conducted in the Langley Unitary Plan wind tunnel. Data were obtained over a Mach number 
range from 1 .60 to 4.65 at Reynolds numbers from approximately 2x10 (exp 6) to 4 x 10(exp 6) per foot depending on the Mach 
number. The data are presented as variations in upwash factor, defined as the slope of the local flow angle to the model angle of 
attack. Some of the effects of yaw angle, longitudinal station, and distance above the surface on the upwash factor are shown. 
Author 
Upwash; Supersonic Speed; Wind Tunnel Tests; Blunt Bodies; Cones; Frustums; Interference Drag; Cylindrical Bodies 

19980227986 NASA Langley Research Center, Hampton, VA USA 

dynamic Characteristics of Ten Booster-Glider Models for Project Dyna-Soar at Mach Numbers of 0.7 



West, F. E., Jr., NASA Langley Research Center, USA; Trescot, Charles D., Jr., NASA Langley Research Center, USA; Wiley, 

Alfred N., Jr., NASA Langley Research Center, USA; 1959; 36p; In English 

Report No.(s): NASA-TM-SX-67; L-647; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A wind-tunnel investigation has been made of ten hypersonic booster-glider models at Mach numbers of 0-7 and 1.0. Only 
the booster portions of the model configurations were varied. Lift, drag, pitching-moment, and base-pressure data were obtained 
for a maximum angle-of-attack range of -10 to 11 deg. These data have not been analyzed since it was desired to expedite the 
publication of the results. 
Author 

Wind Tunnel Tests; Wind Tunnel Models; Hypersonic Gliders; Transonic Speed; Aerodynamic Characteristics; Booster Rocket 
Engines 

19980227087 NASA Ames Research Center, Moffett Field, CA USA 

An Investigation of the Pitching-Moment Contribution of a High Horizontal Tail on an Unswept-Wing and Body Com- 
bination at Mach Numbers from 0.80 to 1.40 

Lippmann, Garth W., NASA Ames Research Center, USA; Aug. 1959; 32p; In English 
Report No.(s): NASA-TM-X-43; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An investigation has been conducted to determine the effects of a high positioned horizontal tail on a wing-body configuration 
having a thin unswept wing of aspect ratio 3.09. Lift and pitching-moment coefficients were obtained for Mach numbers from 
0.80 to 1 .40 at Reynolds numbers of 1 .0 and 1 .5 million and for angles of attack to 20 deg. An experimental study of the pitching- 
moment contribution of the horizontal tail indicated that the marked destabilizing effect of the horizontal tail at high angles of 
attack for Mach numbers of 0.80 to 1 .00 was associated with the formation of completely separated flow on the upper surface of 
the wing. Computations of the interference effects of the wing-body combination on the tail for Mach numbers of 0.80 and 0.94 
and high angles of attack confirmed this conclusion. For a Mach number of 1 .40, and high angles of attack, computations disclosed 
that the destabilizing effect primarily resulted from the trailing vortices of the wing. Two modifications to the basic wing plan 
form, which consisted of chord extensions, were generally unsuccessful in reducing the destabilizing contributions of the horizon- 
tal tail at high angles of attack. 
Author 

Transonic Speed; Unswept Wings; Body-Wing Configurations; Horizontal Tail Surfaces; Aerodynamic Coefficients; Separated. 
Flow; Thin Wings; Lift; Maneuverability ; Pitching Moments; Aircraft Configurations 

19980227089 NASA Ames Research Center, Moffett Field, CA USA 

Transition Reynolds Numbers of Separated Flows at Supersonic Speeds 

Larson, Howard K., NASA Ames Research Center, USA; Keating, Stephen J., Jr., NASA Ames Research Center, USA; Dec. 

1960; 32p; In English 

Report No.(s): NASA-TN-D-349; A-178; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Experimental research has been conducted on the effects of wall cooling, Mach number, and unit Reynolds number on the 
transition Reynolds number of cylindrical separated boundary layers on an ogive-cylinder model. Results were obtained from 

10 



pressure and temperature measurements and shadowgraph observations. The maximum scope of measurements encompassed 
Mach numbers between 2.06 and 4.24, Reynolds numbers (based on length of separation) between 60,000 and 400,000, and ratios 
of wall temperature to adiabatic wall temperature between 0.35 and 1.0. Within the range of tile present tests, the transition 
Reynolds number was observed to decrease with increasing wall cooling, increase with increasing Mach number, and increase 
with increasing unit Reynolds number. The wall cooling effect was found to be four times as great when the attached boundary 
layer upstream of separation was cooled in conjunction with cooling of the separated boundary layer as when only the separated 
boundary layer was cooled. Wall cooling of both the attached and separated flow regions also caused, in some cases, reattachment 
in the otherwise separated region. Cavity resonance present in the separated region for some model configurations was accompa- 
nied by a large decrease in transition Reynolds number at the lower test Mach numbers. 
Author 

Boundary Layer Separation; Separated Flow; Walls; Cooling; Cylindrical Bodies; Ogives; Supersonic Speed; Wall Temperature; 
Reynolds Number 

19980227094 NASA Langley Research Center, Hampton, VA USA 

.4 Wind-Tunnel Investigation of the Development of Lift on Wings in Accelerated Longitudinal Motion 

Turner, Thomas R., NASA Langley Research Center, USA; Aug. 1960; 18p; In English 

Report No.(s): NASA-TN-D-422; L-1027; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An investigation was made in the Langley 300 MPH 7- by 10-foot tunnel to determine the development of lift on a wing during 
a simulated constant-acceleration catapult take-off. The investigation included models of a two-dimensional wing, an unswept 
wing having an aspect ratio of 6, a 35 deg. swept wing having an aspect ratio of 3.05, and a 60 deg. delta wing having an aspect 
ratio of 2.31. All the wings investigated developed at least 90 percent of their steady-state lift in the first 7 chord lengths of travel. 
The development of lift was essentially independent of the acceleration when based on chord lengths traveled, and was in qualita- 
tive agreement with theory. 
Author 

Lift; Delta Wings; Unswept Wings; Swept Wings; Wind Tunnel Tests; Aerodynamic Configurations; Aerodynamic Characteristics; 
Acceleration (Physics) 

19980227096 NASA Ames Research Center, Moffett Field, CA USA 

Large-Scale Wind-Tunnel Teste and Evaluation of the Low-Speed Performance of a 35 deg Sweptback Wing Jet Transport 

Model Equipped with a Blowing Boundary-Layer-Control Flap and Leading-Edge Slat 

Hickey, David H., NASA Ames Research Center, USA; Aoyagi, Kiyoshi, NASA Ames Research Center, USA; Oct. 1960; 54p; 

In English 

Report No.(s): NASA-TN-D-333; A-340; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

A wind-tunnel investigation was conducted to determine the effect of trailing-edge flaps with blowing-type boundary-layer 
control and leading-edge slats on the low-speed performance of a large-scale jet transport model with four engines and a 35 deg. 
sweptback wing of aspect ratio 7. Two spanwise extents and several deflections of the trailing-edge flap were tested. Results were 
obtained with a normal leading-edge and with full-span leading-edge slats. Three-component longitudinal force and moment data 
and boundary-layer-control flow requirements are presented. The test results are analyzed in terms of possible improvements in 
low-speed performance. The effect on performance of the source of boundary-layer-control air flow is considered in the analysis. 
Author 

Boundary Layer Control; Sweptback Wings; Air Flow; Wind Tunnel Tests; Transport Aircraft; Blowing; Externally Blown Flaps; 
Aerodynamic Configurations; Trailing Edge Flaps; Jet Aircraft; Leading Edge Slats 

19980227153 Naval Postgraduate School, Monterey, CA USA 

Supersonic Flow Past Two Oscillating Airfoils 

Alexandris, Georgios, Naval Postgraduate School, USA; Jun. 1998; 85p; In English 

Report No.(s): AD-A350226; No Copyright; Avail: CASI; A05, Hardcopy; A01, Microfiche 

Supersonic flow past two oscillating airfoils is analyzed using an elementary analytical theory valid for low frequencies of 
oscillation. The airfoils may have arbitrary stagger angle. This approach generalizes Sauer's solution for a single airfoil oscillating 
at small frequencies in an unbounded supersonic flow. It is shown that this generalization can provide an elementary theory for 
supersonic flow past two slowly oscillating airfoils. This aerodynamic tool will facilitate the evaluation of pressure distributions 
and consequently the calculation of moment coefficient. Torsional flutter boundaries are computed. The results for the pitch damp- 
ing coefficient are the same when compared with previous analysis. For arbitrary frequencies a linearized method of characteris- 
tics was outlined. The elementary theory that has been developed in the thesis can be used for flutter evaluation of aircraft carrying 

11 



external stores. The result of the thesis is the derivation of the pitch-damping coefficient, which is necessary to predict the flutter 

conditions. 

DTIC 

Supersonic Flow; Airfoils; Oscillations 

19980227178 NASA Langley Research Center, Hampton, VA USA 

Free-Sp inning-Tunnel Investigation of a 1/20-Scale Mode! of the North American T2J-1 Airplane 

Bowman, James S., Jr., NASA Langley Research Center, USA; Healy, Frederick M., NASA Langley Research Center, USA; 

1959; 22p; In English 

Report No.(s): NASA-TM-SX-245; L-872; NASA- AD-3 136; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An investigation has been made in the Langley 20-foot free-spinning tunnel to determine the erect and inverted spin and 
recovery characteristics of a 1/20-scale dynamic model of the North American T2J-1 airplane. The model results indicate that the 
optimum technique for recovery from erect spins of the airplane will be dependent on the distribution of the disposable load. The 
recommended recovery procedure for spins encountered at the flight design gross weight is simultaneous rudder reversal to against 
the spin and aileron movement to with the spin. With full wingtip tanks plus rocket installation and full internal fuel load, rudder 
reversal should be followed by a downward movement of the elevator. For the flight design gross weight plus partially full wingtip 
tanks, recovery should be attempted by simultaneous rudder reversal to against the spin, movement of ailerons to with the spin, 
and ejection of the wing-tip tanks. The optimum recovery technique for airplane-inverted spins is rudder reversal to against the 
spin with the stick maintained longitudinally and laterally neutral. 
Author 
Wind Tunnels; Wind Tunnel Stability Tests; Wind Tunnel Calibration; Scale Models; Loads (Forces) 

19980227180 NASA Ames Research Center, Moffett Field, CA USA 

Evaluation of Blended Wing-Body Combinations with Curved Plan Forms at M ach Numbers Up to 3,50 

Holdaway, George H., NASA Ames Research Center, USA; Mellenthin, Jack A., NASA Ames Research Center, USA; Oct. 1960; 

70p; In English 

Report No.(s): NASA-TM-X-379; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

This investigation is a continuation of the experimental and theoretical evaluation of the effects of wing plan-form variations 
on the aerodynamic performance characteristics of blended wing-body combinations. The present report compares previously 
tested straight-edged delta and arrow models which have leading-edge sweeps of 59.04 and 70-82 deg., respectively, with related 
models which have plan forms with curved leading and trailing edges designed to result in the same average sweeps in each case. 
All the models were symmetrical, without camber, and were generally similar having the same span, length, and aspect ratios. 
The wing sections had an average value of maximum thickness ratio of about 4 percent of the local wing chords in a streamwise 
direction. The wing sections were computed by varying their shapes along with the body radii (blending process) to match the 
selected area distribution and the given plan form. The models were tested with transition fixed at Reynolds numbers of roughly 
4,000,000 to 9,000,000, based on the mean aerodynamic chord of the wing. The characteristic effect of the wing curvature of the 
delta and arrow models was an increase at subsonic and transonic speeds in the lift-curve slopes which was partially reflected in 
increased maximum lift-drag ratios. Curved edges were not evaluated on a diamond plan form because a preliminary investigation 
indicated that the curvature considered would increase the supersonic zero-lift wave drag. However, after the test program was 
completed, a suitable modification for the diamond plan form was discovered. The analysis presented in the appendix indicates 
that large reductions in the zero-lift wave drag would be obtained at supersonic Mach numbers if the leading- and trailing-edge 
sweeps are made to differ by indenting the trailing edge and extending the root of the leading edge. 
Author 

Body-Wing Configurations; Supersonic Speed.; Aerodynamic Characteristics; Wind Tunnel Tests; Planforms; Arrow Wings; Delta 
Wings 

19980227195 NASA Dryden Flight Research Center, Edwards, CA USA 

Preliminary Base Pressures Obtained from the X-I5 Airplane at Mach Numbers from 1 .1. to 3.2 

Saltzman, Edwin J., NASA Dryden Flight Research Center, USA; Aug. 1961; 28p; In English 

Report No.(s): NASA-TN-D-1056; H-215; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Base pressure measurements have been made on the fuselage, 10 deg. -wedge vertical fin, and side fairing of the X-15 airplane. 
Data are presented for Mach numbers between 1.1 and 3.2 for both powered and unpowered flight. Comparisons are made with 
data from small-scale-model tests, semiempirical estimates, and theory. The results of this preliminary study show that operation 
of the interim rocket engines (propellant flow rate approximately 70 lb/sec) reduces the base drag of the X-15 by 25 to 35 percent 

12 



throughout the test Mach number range. Values of base drag coefficient for the side fairing and fuselage obtained from X-15 wind- 
tunnel models were adequate for predicting the overall full-scale performance of the test airplane. The leading-edge sweep of the 
upper movable vertical fin was not an important factor affecting the fin base pressure. The power-off base pressure coefficients 
of the upper movable vertical fin (a 10 deg. wedge with chord-to-thickness ratio of 5.5 and semispan-to-fhickness ratio of 3.2) 
are in general agreement with the small-scale blunt -trailing-edge-wing data of several investigators and with two-dimensional 
theory. 
Author 
Base Pressure; Pressure Measurement; Pressure Ratio; Aerodynamic Coefficients; Transonic Speed; X-15 Aircraft 

19980227199 NASA Langley Research Center, Hampton, VA USA 

Investigation of Interference of a Deflected Jet with Free Stream and Ground on Aerodynamic Characteristics of si Semi- 
spaa Delta- Wing VTOL Model 

Spreemann, Kenneth P., NASA Langley Research Center, USA; Aug. 1961; 92p; In English 
Report No.(s): NASA-TN-D-915; L-1466; No Copyright; Avail: CASI; A05, Hardcopy; A01, Microfiche 

An investigation of the mutual interference effects of the ground, wing, deflected jet stream, and free stream of a semispan 
delta-wing VTOL model at zero and low forward speeds has been conducted in the 17-foot test section of the Langley 300-MPH 
7-by 10-foot tunnel. The model consisted of two interchangeable semispan clipped delta wings, a simplified fuselage, and a high- 
pressure jet for simulation of a jet exhaust. Attached to the wing behind the jet were various sets of vanes for deflecting the jet 
stream to different turning angles. The effect of ground proximity gave the normally expected losses in lift at zero and very low 
forward speeds (up to about 60 or 80 knots for the assumed wing loading of 100 lb/sq ft); at higher forward speeds ground effects 
were favorable. At low forward speeds, out of ground effect, the model encountered large losses in lift and large nose-up pitching 
moments with the model at low angles of attack and the jet deflected 90 deg or 75 deg (the angles required for VTOL performance 
and very low forward speeds). Rotating the model to higher angles of attack and deflecting the jet back to lower angles eliminated 
these losses in lift. Moving the jet rearward with respect to the wing reduced the losses in lift and the nose-up moments at all speeds 
within the range of this investigation. 
Author 
Jet Exhaust; Deflection; Delta Wings; Aerodynamic Characteristics; Angle of Attack 

19980227297 NASA Ames Research Center, Moffett Field, CA USA 

Axial-Force Reduction by Interference Between Jet and Neighboring Afterbody 

Pitts, William C, NASA Ames Research Center, USA; Wiggins, Lyle E., NASA Ames Research Center, USA; Sep. 1960; 108p; 

In English 

Report No.(s): NASA-TN-D-332; No Copyright; Avail: CASI; A06, Hardcopy; A02, Microfiche 

Experimental results are presented for an exploratory investigation of the effectiveness of interference between jet and after- 
body in reducing the axial force on an afterbody with a neighboring jet. In addition to the interference axial force., measurements 
are presented of the interference normal force and the center of pressure of the interference normal force. The free-stream Mach 
number was 2.94, the jet-exit Mach number was 2.71, and the Reynolds number was 0.25 x 10, based on body diameter. The vari- 
ables investigated include static-pressure ratio of the jet (up to 9), nacelle position relative to afterbody, angle of attack (-5 deg 
to 10 deg), and afterbody shape. Two families of afterbody shapes were tested. One family consisted of tangent-ogive bodies of 
revolution with varying length and base areas. The other family was formed by taking a planar slice off a circular cylinder with 
varying angle between the plane and cylinder. The trends with these variables are shown for conditions near maximum jet-after- 
body interference. The interference axial forces are large and favorable. For several configurations the total afterbody axial force 
is reduced to zero by the interference. 
Author 
Afterbodies; Aerodynamic Interference ; Jet Flow; Center of Pressure; Static Pressure; Pressure Ratio 

19980227277 NASA Langley Research Center, Hampton, VA USA 

A Transonic Investigation of Changing Indentation Design Mach Number on the Aeodynamic Characteristics of a 45 deg 

Sweptback-Wing-Body Combination Designed for High Performance 

Loving, Donald L., NASA Langley Research Center, USA; Oct. 1961; 90p; In English 

Report No.(s): NASA-TN-D-941; L-1698; No Copyright; Avail: CASI; A05, Hardcopy; A01, Microfiche 

The effects of changing indentation design Mach number on the aerodynamic characteristics of a 45 deg. sweptback-wing- 
body combination designed for high performance have been investigated at Mach numbers from 0.80 to 1.13 in the Langley 8-foot 
transonic tunnel and at a Mach number of 1.43 in the Langley 8-foot transonic pressure tunnel. The Reynolds number of the inves- 

13 



tigation covered the range from approximately 2.5 x 10 (exp 6) to approximately 3.0 x 10(exp 6) based on the mean aerodynamic 
chord of the wing. The 45 deg. sweptback wing with camber and a thickened root was tested at deg. angle of incidence on an 
unindented body and on bodies indented for Mach numbers M of 1.0, 1.2, and 1.4. Transonic and supersonic area rules were used 
in the design of the indented bodies. Theoretical zero-lift wave drag was calculated for these wing-body combinations. A -2 deg. 
angle of incidence of the wing, and M = 1 .4 revised body indentation, and fixed transition also were investigated. Experimental 
values of zero-lift wave drag for the indented-body combinations followed closely the area-rule concept in that the lowest zero-lift 
wave-drag coefficient was obtained at or near the Mach number for which the body of the combination was designed. Theoretical 
values of zero-lift wave drag were considered to be in good agreement with the experimental results. At a given supersonic Mach 
number the highest values of maximum lift-drag ratio for the various combinations also were obtained at or near the Mach number 
for which the body of the combination was designed. At Mach numbers of 1.0, 1.2, and 1.43, the maximum lift-drag ratios were 
15.3, 13.0, and 9.2, respectively. The use of an angle of incidence of -2 deg. for the wing in combination with the M = 1.2 body 
increased the zero-lift wave drag and decreased the maximum lift-drag ratio. All configurations maintained stable characteristics 
up to the highest lift coefficient of the investigation (C(L) approx. equal to 0.5). 
Author 

Aerodynamic Characteristics; Aerodynamic Coefficients; Sweptback Wings; Wind Tunnel Tests; Body-Wing Configurations; 
Transonic Speed 

19981)227281 NASA Ames Research Center, Moffett Field, CA USA 

Theoretical Pressure Distributions on Wings of Finite Span at Zero Incidence for Mach Numbers Near i 

Alksne, Alberta Y., NASA Ames Research Center, USA; Spreiter, John R., NASA Ames Research Center, USA; 1961; 50p; In 

English 

Report No.(s): NASA-TR-R-88; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A method employed heretofore by the authors to obtain approximate solutions of the transonic flow equation for plane and 
axisymmetric flow is extended to give reasonable result for wings of finite span, consistent with the known properties of transonic 
flows. In this method the partial differential equation appropriate to the study of transonic flow is replaced by a nonlinear ordinary 
differential equation which can be solved by numerical methods, Asymptotic forms of this differential equation are given for very 
high and very low aspect ratios and analytic results are obtained for certain special cases. Numerical results, calculated by use 
of electronic computing machines, are given in the form of pressure distribution and pressure drag for two profile shapes, wedge 
and circular are, for wings of rectangular plan form. The range of aspect ratios covered extends effectively from zero to infinity 
and agreement with asymptotic results is shown at both limits. 
Author 

Pressure Distribution; Wings; Pressure Drag; Axisymmetric Flow; Approximation; Numerical, Analysis; Low Aspect Ratio; Tran- 
sonic Flow 

19980227287 NASA Ames Research Center, Moffett Field, CA USA 

The Use of Drag Modulation to Limit the Mate at Which Deceleration Increases Daring Nonlifting Entry 

Levy, Lionel L., Jr., NASA Ames Research Center, USA; Sep. 1961; 32p; In English 

Report No.(s): NASA-TN-D-1037; A502; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The method developed in NASA TN D-319 for studying the atmosphere entry of vehicles with varying aerodynamic forces 
has been applied to obtain a closed-form solution for the motion, heating, range, and variation of the vehicle parameter m/C(D)A 
for nonlifting entries during which the rate of increase of deceleration is limited. The solution is applicable to vehicles of arbitrary 
weight, size, and shape, and to arbitrary atmospheres. Results have been obtained for entries into the earth's atmosphere at escape 
velocity during which the maximum deceleration and the rate at which deceleration increases were limited. A comparison of these 
results with those of NASA TN D-319, in which only the maximum deceleration was limited, indicates that for a given corridor 
depth, limiting the rate of increase of deceleration and the maximum deceleration requires an increase in the magnitude of the 
change in M/C(D)A and results in increases in maximum heating rate, total heat absorbed at the stagnation point, and range. 
Author 
Atmospheric Entry; Deceleration; Drag; Aerodynamic Forces; Aerospace Vehicles; Aeromaneuvering 

19980227303 NASA Ames Research Center, Moffett Field, CA USA 

Wind Tunnel Investigation at Supersonic Speeds of the Lift, Drag, Static-Stability and Control Characteristics of a 

0.03-Scale Model of an Interim Development Version of the B-70 Airplane 

Daugherty, James C, NASA Ames Research Center, USA; Green, Kendal H., NASA Ames Research Center, USA; Jun. 07, 1961; 

154p; In English 

14 



Report No.(s): NASA-TM-SX-572; AF-AM-199; A-407; No Copyright; Avail: CASI; A08, Hardcopy; A02, Microfiche 

A 0.03-scale model of an interim development version of the B-70 airplane was tested at Mach numbers of 2.5, 3.0, and 3.5 
at a Reynolds number of 5 million. Canard, elevon, aileron, and rudder control effectiveness were measured. A complete compo- 
nent drag evaluation study was made. Effects of modifications of the forebody boundary-layer gutter and inlet cowl-lip angle were 
obtained. In addition, the effects of increased wing camber were determined. Tests were conducted on the basic configuration with 
both natural and fixed boundary-layer transition. Trim drag estimates are made and the data are extrapolated to flight conditions 
at a Mach number of 3.0 to give maximum trimmed lift-drag ratios. 
Author 
Lift Drag Ratio; Boundary Layer Transition; Supersonic Speed; Wind Tunnel Tests; Aerodynamic Drag; Rudders; Ailerons 

19980227395 NASA Ames Research Center, Moffett Field, CA USA 

The Shock-Wave Patterns on a Cranked- Wing Configuration 

Sammonds, Robert I., NASA Ames Research Center, USA; Nov. 1960; 16p; In English 

Report No.(s): NASA-TN-D-346; A-433; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The shock- wave patterns of a complex configuration with cranked cruciform wings and a cone-cylinder body were examined 
to determine the interaction of the body bow wave with the flow field about the wing. Also of interest, was the interaction of the 
forward (760 sweptback) wing leading-edge wave with the rear (600 sweptback) wing leading-edge wave. The shadowgraph pic- 
tures of the model in free flight at a Mach number of 4.9, although not definitive, appear to indicate that the body bow wave crosses 
the outer wing panel after first being refracted either by the leading-edge wave of the 600 sweptback wing or by pressure fields 
in the flow crossing the wing. 
Author 

Shock Waves; Cruciform, Wings; Bow Waves; Wing Planforms; Sweptback Wings; Aerodynamic Characteristics; Flow 
Distribution 

19980227361 NASA Langley Research Center, Hampton, VA USA 

Heat Transfer to 36.75 and 45 degree Swept Blunt Leading Edges in Free Flight at Mach Numbers from 1.70 to 2.99 and 



ONeal, Robert L., NASA Langley Research Center, USA; Mar. 1960; 40p; In English 

Report No.(s): NASA-TM-X-208; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A flight investigation has been conducted to study the heat transfer to swept-wing leading edges. A rocket-powered model 
was used for the investigation and provided data for Mach number ranges of 1.78 to 2.99 and 2.50 to 4.05 with corresponding 
free-stream Reynolds number per foot ranges of 13.32 x 10(exp 6) to 19.90 x 10(exp 6) and 2.85 x 10(exp 6) to 4.55 x 10(exp 
6). The leading edges employed were cylindrically blunted wedges ', three of which were swept 450 with leading-edge diameters 
of 1/4, 1/2, and 3/4 inch and one swept 36-750 with a leading-edge diameter of 1/2 inch. In the high Reynolds number range, mea- 
sured values of heat transfer were found to be much higher than those predicted by laminar theory and at the larger values of lead- 
ing-edge diameter were approaching the values predicted by turbulent theory. For the low Reynolds number range a comparison 
between measured and theoretical heat transfer showed that increasing the leading-edge diameter resulted in turbulent flow on 
the cylindrical portion of the leading edge. 
Author 
Heat Transfer; Swept Wings; Blunt Leading Edges; Free Flow; Turbulent Flow; Cylindrical Bodies 

19980227409 NASA Ames Research Center, Moffett Field, CA USA 

adinal Force and Moment Data at Mach Numbers from 0,60 to 1.40 for a Family of Elliptic Cones with Various 
isex Angles 

Stivers, Louis S., Jr., NASA Ames Research Center, USA; Levy, Kionel L., Jr., NASA Ames Research Center, USA; Dec. 1961; 
38p; In English 
Report No.(s): NASA-TN-D-1149; A-548; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An investigation has been made to determine the aerodynamic characteristics of four elliptic cones having plan-form semi- 
apex angles ranging from about 9 to 31 deg., and also for one of these cones modified on the upper surface to reduce the base area 
by about one half. The tests were made for angles of attack from about -2 to +21 deg., at Mach numbers from 0.60 to 1.40, and 
for a constant Reynolds number of 1.4 million, based on the length of the models. For each model, lift, pitching-moment, and drag 
coefficients, and lift-drag ratios are presented for the forebody, and axial-force coefficients are presented for the base. Calculated 
lift and pitching- moment curves for the elliptic cones, and lift-curve slopes for each model at supersonic Mach numbers are shown 

15 



for comparison with the corresponding experimental values. Lift-drag ratios are also given for the forebody and base combined. 

These data are presented without discussion. 

Author 

Aerodynamic Characteristics; Aerodynamic Coefficients; Aerodynamic Forces; Moment Distribution; Cones; Transonic Speed; 
Forebodies; Aerodynamic Configurations 

19980227411 NASA Langley Research Center, Hampton, VA USA 

Approximate Temperature Distributions and Streamwise Heat Conduction Effects in the Transient Aerodynamic Healing 

of Thin-Skinned Bodies 

Conti, Raul J., NASA Langley Research Center, USA; Sep. 1961; 92p; In English 

Report No.(s): NASA-TN-D-895; L-1227; No Copyright; Avail: CASI; A05, Hardcopy; A01, Microfiche 

An approximate method is devised to determine temperature distributions during the transient aerodynamic heating of thin- 
skinned, heat-conducting bodies. This permits evaluation of the streamwise conduction errors arising in the measurement of heat- 
transfer coefficients based on the skin-temperature history. The present method is valid for a large range of body shapes and 
thickness distributions, within the limitations of one-dimensional (streamwise) heat conduction, quasi-isothermal surface, 
constant adiabatic wall temperature, and negligible radiative heat transfer. Numerical computations were carried out for flat plates, 
wedges, and conical, hemispherical, and hemicylindrical shells. The results are presented in the form of nondimensional charts 
that permit a rapid evaluation of a 10-percent error threshold in transient heat-transfer measurements. 
Author 

Conductive Heat Transfer; Aerodynamic Heating; Surface Temperature; Transient Heating; Conical Shells; Flat Plates; Wedges; 
Hemispherical Shells 

1998022741.4 NASA Ames Research Center, Moffett Field, CA USA 

Exploratory Study of the Reduction in Friction Drag Due to Streamwise Injection of Helium 

Swenson, Byron L., NASA Ames Research Center, USA; Jan. 1961; 34p; In English 

Report No.(s): NASA-TN-D-342; A-414; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The effects on average skin-friction drag and pressure drag of the streamwise injection of helium into the boundary layer near 
the nose of a 6 deg. half-angle cone at Mach numbers of 3 to 5 are presented. Large reductions in skin friction are shown to be 
possible with relatively small amounts of helium injection. 
Author (revised) 

Skin Friction; Pressure Drag; Half Cones; Wind Tunnel Tests; Gas Injection; Helium; Supersonic Speed; Friction Drag; Drag 
Reduction 

19980227417 NASA Langley Research Center, Hampton, VA USA 

Spin-Tunnel Investigation of a 1/20-Scale Model of the Northrop F-5E Airplane 

Scher, Stanley H., NASA Langley Research Center, USA; White, William L., NASA Langley Research Center, USA; Sep. 1977; 

48p; In English 

Contract(s)/Grant(s): RTOP 505-11-41-08 

Report No.(s): NASA-TM-SX-3556; L-11541; AF-AM-422; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An investigation has been conducted in the Langley spin tunnel to determine the spin and recovery characteristics of a 
1/20-scale model of the Northrop F-5E airplane. The investigation included erect and inverted spins, a range of center-of- gravity 
locations and moments of inertia, symmetric and asymmetric store loadings, and a determination of the parachute size required 
for emergency spin recovery. The effects of increased elevator trailing-edge-up deflections, of leading-edge and trailing-edge flap 
deflections, and of simulating the geometry of large external stores were also determined. 
Author 
F-5 Aircraft; Scale Models; Wind Tunnel Tests; Aircraft Spin; Spin Dynamics; Control Stability; Aircraft Control 

19980227422 Institute of Theoretical and Applied Mechanics, Novosibirsk, Russia 

International Conference on the Methods of Aerophysical Research 1998 "ICMAR 98", Part 1 

1998; 259p; In English; Methods of Aerophysical Research, 29 Jun. - 3 Jul. 1998, Novosibirsk, Russia 

Contract(s)/Grant(s):F61775-98-WE002 

Report No.(s): AD-A350416; EOARD-CSP-98-1025; No Copyright; Avail: CASI; A12, Hardcopy; A03, Microfiche 

16 



The Final Proceedings for International Conference on Methods of Aerophysical Research (ICMAR'98), 29 June 1998 - 3 
July 1998 This is an interdisciplinary conference. Topics include: Problems of Modeling at sub/trans/super/hypersonic velocities; 
Methods of flow diagnostics; Instrumentation for aerophysical experiments; Verification of CFD models and methods. 
DTIC 

Conferences; Computational Fluid Dynamics; Atmospheric Physics 

19980227430 NASA Langley Research Center, Hampton, VA USA 

Subsonic Aerodynamic Characteristics of the Mil 7 Bomb with a Fragmentation Wrap 

Capone, Francis J., NASA Langley Research Center, USA; Jul. 1965; 32p; In English 

Report No.(s): NASA-TM-SX-1106; A-AM-77; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The aerodynamic effects of a fragmentation wrap around the cylindrical section of the Ml 17 bomb have been determined 
in the Langley 16-foot transonic tunnel at Mach numbers from 0.30 to 0.90 and angles of attack from approximately -4 to 8 deg. 
The bomb without the wrap was also tested to a Mach number of 1.15. Total and static pressures measured at a fuze arming mecha- 
nism are also presented. The test Reynolds number based on a model length of 89.83 inches (228.17 cm) varied from 14.29 x 
10(exp 6) to 29.99 x 10(exp 6). 
Author 

Aerodynamic Characteristics; Subsonic Speed; Bombs (Ordnance); Wind Tunnel Tests; Cylindrical Bodies; Aerodynamic 
Coefficients 

19980227736 NASA Langley Research Center, Hampton, VA USA 

Study of Flow Over Oscillating Airfoil Models sit a Mach Number of 7,0 in Helium 

Arman, Ali, NASA Langley Research Center, USA; Dec. 1961; 24p; In English 

Report No.(s): NASA-TN-D-992; L-839; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A wind-tunnel study of unsteady flow at a Mach number of 7 in helium has been conducted on several sting-mounted wedge, 
double-wedge, and flat-plate airfoil models with three different leading-edge radii. The data were obtained by taking high-speed 
schlieren motion pictures of the decaying motion of the model as it was released from an initial deflection. The shock- wave posi- 
tion observed on the sharp-leading-edge models during the oscillation was compared with that obtained by use of unsteady flow 
theory as well as steady-state theory. Comparison of theoretical results indicated that no unsteady-flow effects exist over the range 
of reduced frequencies k, 0.007 less than equal than k less than or equal 0.030, studied experimentally. The experimental results 
confirmed this finding as no unsteady-flow effects were detected in this reduced-frequency range. Comparison of shock-wave 
positions measured for the blunt models with those calculated by steady-state methods indicated fair agreement. 
Author 
Unsteady Flow; Wind Tunnel Tests; Airfoil Oscillations; Aeroelasticity; Hypersonic Speed; Flat Plates; Leading Edges 

19980227752 NASA Langley Research Center, Hampton, VA USA 

Equations for the Induced Velocities Near a Lifting Motor with Nonuniform Azimuthwise Vorticity Distribution 

Heyson, Harry H., NASA Langley Research Center, USA; Aug. 1960; 28p; In English 

Report No.(s): NASA-TN-D-394; L-797; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Equations, which can be integrated on high-speed computing machines, are developed for all three components of induced 
velocity at an arbitrary point near the rotor and for an arbitrary harmonic variation of vorticity. Sample calculations for vorticity 
which varies as the sine of the azimuth angle indicate that the normal component of induced velocity is, in this case, uniform along 
either side of the lateral axis. 
Author 

Rotor Aerodynamics; Lifting Rotors; Vorticity; Flow Distribution; Flow Velocity; Velocity Distribution; Aerodynamic 
Interference 

19980227753 NASA Langley Research Center, Hampton, VA USA 

Free-Spinning-Tunnel Investigation of a 1/20-Scale Model of an Unswept- Wing Jet-Propelled Trainer Airplane 

Bowman, James S., Jr., NASA Langley Research Center, USA; Healy, Frederick M., NASA Langley Research Center, USA; Jun. 

1960; 20p; In English 

Report No.(s): NASA-TN-D-381; L-872; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A flutter analysis employing the kernel function for three- dimensional, subsonic, compressible flow is applied to a flutter- 
tested tail surface which has an aspect ratio of 3.5, a taper ratio of 0.15, and a leading-edge sweep of 30 deg. Theoretical and experi- 
mental results are compared at Mach numbers from 0.75 to 0.98. Good agreement between theoretical and experimental flutter 

17 



dynamic pressures and frequencies is achieved at Mach numbers to 0.92. At Mach numbers from 0.92 to 0.98, however, a second 
solution to the flutter determinant results in a spurious theoretical flutter boundary which is at a much lower dynamic pressure 
and at a much higher frequency than the experimental boundary. 
Author 

Wind Tunnel Tests; Unswe.pt Wings; Aircraft Spin; Subsonic Flow; Scale Models; Flutter Analysis; Spin Dynamics 

19981)227791 NASA Ames Research Center, Moffett Field, CA USA 

The Numerical Calculation of Flow Past Conical Bodies Supporting Elliptic Conical Shock Waves at Finite Angles of Inci- 
dence 

Briggs, Benjamin R., NASA Ames Research Center, USA; Nov. 1960; 68p; In English 
Report No.(s): NASA-TN-D-340; A-385; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

The inverse method, with the shock wave prescribed to be an elliptic cone at a finite angle of incidence, is applied to calculate 
numerically the supersonic perfect-gas flow past conical bodies not having axial symmetry. Two formulations of the problem are 
employed, one using a pair of stream functions and the other involving entropy and components of velocity. A number of solutions 
are presented, illustrating the numerical methods employed, and showing the effects of moderate variation of the initial 
parameters. 
Author 
Conical Bodies; Shock Waves; Supersonic Flow; Aerodynamic Configurations; Ideal Gas 

19980227792 NASA Langley Research Center, Hampton, VA USA 

Transonic Wind-Tunnel Investigation of the Fin Loads on a 1/8-Scale Mode! Simulating the First Stage of the Scout 

Research Vehicle 

Kelly, Thomas C, NASA Langley Research Center, USA; Jun. 1961; 50p; In English 

Report No.(s): NASA-TN-D-918; L-1438; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An investigation to determine the fin loads on a 1/8-scale model simulating the first stage of the Scout research vehicle was 
made in the Langley 8-foot transonic tunnel at Mach numbers from 0.40 to 1.20. Tests were conducted over an angle-of-attack 
range from about -10 to 10 deg and at a Reynolds number per foot of approximately 3.5 x 10(exp 6). Results of the tests indicate 
that for a given angle of attack, negative tip-control deflections caused decreases in normal-force and fin-bending-moment coeffi- 
cients and increases in pitching-moment coefficient, as would be expected. The effects were slight at a model angle of attack of 
-10 deg where tip-control stall had probably occurred but increased with an increase in angle of attack. 
Author 
Aerodynamic Loads; Wind Tunnel Tests; Transonic Speed; Fins; Research Vehicles; Scale Models; Aerodynamic Coefficients 

19980227793 NASA Langley Research Center, Hampton, VA USA 

Wind-Tunnel Investigation of a Balloon as a Towed Decelerates at Mach Numbers from 1.47 to 2.50 

McShera, John T, NASA Langley Research Center, USA; Keyes, J. Wayne, NASA Langley Research Center, USA; Aug. 1961; 

82p; In English 

Report No.(s): NASA-TN-D-919; L-884; No Copyright; Avail: CASI; A05, Hardcopy; A01, Microfiche 

A wind-tunnel investigation has been conducted to study the characteristics of a towed spherical balloon as a drag device at 
Mach numbers from 1.47 to 2.50, Reynolds numbers from 0.36 x 10(exp 6) to 1.0 x 10(exp 6) , and angles of attack from -15 to 
15 deg. Towed spherical balloons were found to be stable at supersonic speeds. The drag coefficient of the balloon is reduced by 
the presence of a tow cable and a further reduction occurs with the addition of a payload. The balloon inflation pressure required 
to maintain an almost spherical shape is about equal to the free-stream dynamic pressure. Measured pressure and temperature dis- 
tribution around the balloon alone were in fair agreement with predicted values. There was a pronounced decrease in the pressure 
coefficients on the balloon when attached to a tow cable behind a payload. 
Author 
Balloons; Supersonic Speed; Towed Bodies; Wind Tunnel Tests; Aerodynamic Coefficients; Free Flow 

19980227802 NASA Langley Research Center, Hampton, VA USA 

Aerodynamic Characteristics, Temperature, and Noise Measurements of a Large-Scale External-Flow Jet-A 

Flap Model with Turbojet Engines Operating 

Fink, Marvin P., NASA Langley Research Center, USA; Sep. 1961; 50p; In English 

Report No.(s): NASA-TN-D-943; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

18 



An investigation has been conducted in the Langley full-scale tunnel on a large-scale model powered by turbojet engines with 
flattened rectangular nozzles. The wing had 35 deg. sweep of the leading edge, an aspect ratio of 6.5, a taper ratio of 0.3 1, and 
NACA 65(1)-412 and 65-408 airfoils at the root and tip. The investigation included measurements of the longitudinal aerody- 
namic characteristics of the model with half-span and full-span flaps and measurements of the sound pressure and skin temperature 
on the portions of the lower surface of the wing immersed in the jet flow. The tests were conducted over a range or angles of attack 
from -8 to 16 deg. for Reynolds numbers from 1.8 x 10(exp 6) to 4.4 x 10(exp 6) and a range of momentum coefficients from 
to 2.0. In general, the aerodynamic results of this investigation made with a large-scale hot-jet model verified the results of pre- 
vious investigations with small models powered by compressed-air jets. Although blowing was only done over the inboard portion 
of the wing, substantial amounts of induced lift were also obtained over the outboard portion of the wing. Skin temperatures were 
about 340 F and wing heating could be handled with available materials without cooling. Random acoustic loadings on the wing 
surface were high enough to indicate that fatigue failure from this source would require special consideration in the design of an 
external-flow jet flap system for an airplane. 
Author 

Aerodynamic Characteristics; Noise Measurement; Temperature; Turbojet Engines; Airfoils; Wind Tunnel Models; Scale Mod- 
els; Wind Tunnel Tests; Lift Augmentation; Sweptback Wings; Jet Flow 

19980227803 NASA Langley Research Center, Hampton, VA USA 

Experimental and Theoretical Deflections and Natural Frequencies of an Inflatable Fabric Plate 

Stroud, W. Jefferson, NASA Langley Research Center, USA; Oct. 1961; 32p; In English 

Report No.(s): NASA-TN-D-931; L-1317; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Static and vibration tests were performed on an inflatable square fabric plate supported on all edges. Lateral deflections and 
natural frequencies showed good agreement with calculations made using a linear small-deflection theory. 
Author 

Structural Analysis; Inflatable Structures; Static Tests; Vibration Tests; Plates (Structural Members); Fabrics; Deflection; Reso- 
nant Frequencies 



03 
AIR TRANSPORTATION AND SAFETY 

Includes passenger and cargo air transport operations; and aircraft accidents. 

19980221785 European Organization for the Safety of Air Navigation, Bretigny-sur-Orge, France 

ATFM Studies; Remaining Overdeliveries 

Ganvert, E., European Organization for the Safety of Air Navigation, France; Greiling, Y., European Organization for the Safety 

of Air Navigation, France; Vidal, A., European Organization for the Safety of Air Navigation, France; Mar. 1998; 26p; In English 

Report No.(s): PB98-164304; EEC/NOTE-5/98; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

This document describes an ATFM (Air Traffic Flow Management) study conducted by the Centre of Expertise Flight Data 
Research on behalf of the CFMU (Central How Management Unit) in to evaluate the performance of the current CFMU Opera- 
tions and to evaluate the Slot Allocation process by analyzing the remaining overdeliveries on regulated sectors. 
NTIS 
Air Traffic Control; Airports; Flight Management Systems; Flow Distribution 

19980221788 European Organization for the Safety of Air Navigation, Experimental Centre, Bretigny-sur-Orge, France 

Coverage of European Air Traffic for the Base Aircraft Data (BADA) 

Bos, A., European Organization for the Safety of Air Navigation, France; Mar. 1998; 30p; In English 

Report No.(s): PB98-164320; EEC/NOTE-8/98-Rev-3.0; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The air traffic statistics from the CFMU for December 1997 and January 1998 are used to determine the coverage of European 
air traffic by the Base of Aircraft Data (BADA) Revision 3.0 BADA consists of a set of aircraft models used at the EEC and other 
European research institutes for aircraft trajectory simulation. The results show that the 67 aircraft types within BADA 3.0 cover 
89.4% of the European air traffic. The addition of 1 type would bring the coverage to the target of 90%. 
NTIS 
Air Traffic; Research Aircraft; Aircraft Models 

19 



19980221792 Federal Aviation Administration, Washington, DC USA 

Notices to Airmen; Domestic/International 

Jul. 16, 1998; 238p; In English 

Report No.(s): PB98-163389; No Copyright; Avail: CASI; All, Hardcopy; A03, Microfiche 

Table of Contents: Airway Notams; Airports, Facilities, and Procedural Notams; General FDC Notams; Part 95 Revisions 
to Minimum En Route IFR Altitudes and Changeover Points; International Notices to Airmen; and Graphic Notices. 
NTIS 
National Airspace System; Air Navigation; Airports; Altitude; Graphs (Charts); Constrictions 

19980221806 Federal Aviation Administration, FAA Technical Center, Atlantic City, NJ USA 

Functional Requirements for Screener Assist Technologies 

Fobes, J. L., Federal Aviation Administration, USA; Neiderman, Eric C, Federal Aviation Administration, USA; Jul. 1998; 36p; 

In English 

Report No.(s): PB98-159742; DOT/FAA/AR-98/35; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

This document lists the human factors functional requirements for Screener Assist Technologies (SAT) to enhance screener 
performance to detect threat objects. The report also describes the required interactions with Threat Image Projection (TIP) sys- 
tems, naming conventions for threats, data report capabilities, FAA acceptance test procedures, and operational and technical cri- 
teria that will be used to assess system effectiveness. 
NTIS 
Functional Design Specifications; Human Factors Engineering; Aircraft Safety; Airport Security; X Ray Inspection 

19980223924 Texas Univ., Health Science Center, Houston, TX USA 

General Aviation Accidents; The USA Air Force Aero Club Solution 

Brandt, Keith E., Texas Univ., USA; Aug. 07, 1998; 72p; In English 

Report No.(s): AD-A350974; AFIT-98-049; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

Aviation is an intrinsically safe mode of travel. In 1994, the United States Air force system of Aero Clubs put forth substantial 
effort to put a program in place (fly Smart) to improve flying safety in its aircraft. This study compares the accident rates of Aero 
Club aircraft with rates seen in general aviation. A comparison is also made of the years prior to implementation of fly Smart to 
the three years following implementation. Aero Club records of accidents were available from 1987 through 1997. General avi- 
ation mishap statistics are collected by the National Transportation Safety Board and are collected and presented to the public by 
the Aircraft Owners and Pilots Association in the form of an annual general aviation report. Comparison of these figures show 
that the Aero Club system had a lower accident rate and fatality rate in all but one study year (1992, Aero Club 10.12 accidents 
and 2.38 fatal accidents per 100,000 flying hours; general aviation 8.97 accidents and 1.75 fatal accidents per 100,000 flying 
hours). The Aero Club accident rate in the period following implementation of fly Smart (1995 - 1997) was lower than before 
implementation (1987 - 1993, 5.19 versus 1.63, p=0.047), while general aviation rates for the same periods were unchanged (8.29 
versus 8.00, p>0.05). No differences were seen in rates of larger vs. mid-size or small clubs. There were no differences in the acci- 
dent rates of closed vs. open clubs. The Air Force Aero Clubs are certainly more restrictive than general aviation, but the improve- 
ment in safety record suggests the tighter regulations are rules you can live with. 
DTIC 
Aircraft Accidents; Flight Safety; Safety Management; Civil Aviation 

19980227159 Naval Postgraduate School, Monterey, CA USA 

Allocating Flight Hours to Army Helicopters 

Pippin, Bradley W., Naval Postgraduate School, USA; Jun. 1998; 55p; In English 

Report No.(s): AD-A350138; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

Army helicopter battalions, consisting of 24 helicopters valued from $206.4 million (UH-60 Blackhawk battalion) to $432 
million (AH-64 Apache battalion), allocate flight hours to helicopters using manual techniques that have caused an unnecessary 
decrease in battalion deployability. This thesis models the battalion's flight hour allocation problem using optimization; it devel- 
ops both a mixed integer linear program and a quadratic program. The 2nd Battalion, 4th Aviation Regiment of 4th Mechanized 
Division currently uses a spreadsheet implementation of the quadratic program developed by the author called QFHAM (Qua- 
dratic Flight Hour Allocation Model), that is available to other battalions for use with existing software and computer resources. 
The mixed integer linear program, called FHAM (Flight Hour Allocation Model) more appropriately models the problem, but 
requires additional software. This thesis validates the two models using actual flight hour data from a UH-60 battalion under both 
typical training and contingency scenarios. The models provide a monthly flight hour allocation for the battalion's aircraft that 

20 



results in a steady-state sequencing of aircraft into phase maintenance, thus eliminating phase maintenance backlog and providing 
a fixed number of aircraft available for deployment. This thesis also addresses the negative impact of current helicopter battalion 
readiness measures on deployment and offers alternatives. 
DTIC 

Scheduling; Allocations; AH-64 Helicopter; Computer Systems Programs 

19980227419 NASA Lewis Research Center, Cleveland, OH USA 

.4 Combined Water-Bromotritluoromcthane Crash-Fire Protection System for a T-56 Turbopropeller Engine 

Campbell, John A., NASA Lewis Research Center, USA; Busch, Arthur M., NASA Lewis Research Center, USA; Aug. 1959; 

36p; In English 

Report No.(s): NASA-TN-D-28; E-308; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A crash-fire protection system is described which will suppress the ignition of crash-spilled fuel that may be ingested by a 
T-56 turbo-propeller engine. This system includes means for rapidly extinguishing the combustor flame, means for cooling and 
inerting with water the hot engine parts likely to ignite engine ingested fuel, and means for blanketing with bromotrifluoromethane 
massive metal parts that may reheat after the engine stops rotating. Combustion-chamber flames were rapidly extinguished at the 
engine fuel nozzles by a fuel shutoff and drain valve. Hot engine parts were inerted and cooled by 42 pounds of water discharged 
at seven engine stations. Massive metal parts that could reheat were inerted with 10 pounds of bromotrifluoromethane discharged 
at two engine stations. Performance trials of the crash-fire protection system were conducted by bringing the engine up to takeoff 
temperature, actuating the crash-fire protection system, and then spraying fuel into the engine to simulate crash-ingested fuel. No 
fires occurred during these trials, although fuel was sprayed into the engine from 0.3 second to 15 minutes after actuating the crash- 
fire protection system. 
Author 

Fire Prevention; Engine Parts; T-56 Engine; Extinguishing; Fire Extinguishers; Crashes; Aircraft Fuels; Spilling; Spraying; 
Water 

04 
AIRCRAFT COMMUNICATIONS AND NAVIGATION 

Includes digital and voice communication with aircraft; air navigation systems (satellite and ground based); and air traffic control. 

19980223080 Rockwell Collins, Inc., Advanced Technology Center, Cedar Rapids, IA USA 

Integrated Airport Surface Operations 

Koczo, S., Rockwell Collins, Inc., USA; Jul. 1998; 180p; In English 

Contract(s)/Grant(s): NAS 1-19704; RTOP 538-04-13-02 

Report No.(s): NASA/CR-1998-208441; NAS 1.26:208441; No Copyright; Avail: CASI; A09, Hardcopy; A02, Microfiche 

The current air traffic environment in airport terminal areas experiences substantial delays when weather conditions deterio- 
rate to Instrument Meteorological Conditions (IMC). Research activity at NASA has culminated in the development, flight test 
and demonstration of a prototype Low Visibility Landing and Surface Operations (LVLASO) system. A NASA led industry team 
and the FAA developed the system which integrated airport surface surveillance systems, aeronautical data links, DGPS naviga- 
tion, automation systems, and controller and flight deck displays. The LVLASO system was demonstrated at the Hartsfield- 
Atlanta International Airport using a Boeing 757-200 aircraft during August, 1997. This report documents the contractors role 
in this testing particularly in the area of data link and DGPS navigation. 
Author 

Air Traffic Control; Data Links; All-Weather Landing Systems; Boeing 757 Aircraft; Instrument Flight Rules; All-Weather Air 
Navigation; Navigation Instruments 

19980227146 Federal Aviation Administration, Technical Center, Atlantic City, NJ USA 

Traffic Information Service (ITS) Developmental/Operational Test and Evaluation (DT/E and OT/E) Final Report 

McNeil, Michael, Federal Aviation Administration, USA; Sharkey, Robert, Federal Aviation Administration, USA; Jun. 1998; 

145p; In English 

Report No.(s): AD-A350376; DOT/FAA/CT-TN98/10; No Copyright; Avail: CASI; A07, Hardcopy; A02, Microfiche 

The Federal Aviation Administration (FAA) Traffic Information Service (TIS) Developmental Test and Evaluation (DT&E) 
and Operational Test and Evaluation (OT&E) Final Test Report is prepared by the Mode Select (Mode S) Test Group of the Sur- 
veillance Branch ACT-310. It provides the detailed analysis, results, the final conclusions, and recommendations drawn from the 

21 



DT&E and OT&E of the TIS data link service for the Mode S Beacon Radar System. The purpose of the TIS data link function 
is intended to improve the safety and efficiency of "see-and-avoid" flight by providing automatic display to the pilot of nearby 
traffic and warnings of any potentially threatening conditions. The source of TIS information is the file of aircraft tracks main- 
tained by the ground Mode S sensor providing coverage for a region of airspace. 
DTIC 
Air Traffic; Systems Analysis; Data Links; Beacons; Airspace 

19980227148 Federal Aviation Administration, Technical Center, Atlantic City, NJ USA 

Reduced Horizontal Separation Minima (RHSM) Concept Exploration Simulation 

Elkan, Elizabeth, Federal Aviation Administration, USA; Kopardekar, Parimal, Federal Aviation Administration, USA; Stahl, 

David, Federal Aviation Administration, USA; Mar. 1998; 44p; In English 

Report No.(s): AD-A350324; DOT/FAA/CT-TN97/3; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The Informal South Pacific Air Traffic Services Coordinating Group has been investigating a number of concepts to improve 
operational efficiency for flights in the Pacific Oceanic region. The Federal Aviation Administration (FAA) Air Traffic Require- 
ments (ATR-3 10) and Air Traffic Operations (ATO-100) program offices tasked the Simulation and Systems Integration Branch 
(ACT-540), in cooperation with the Oceanic and Offshore Integrated Product Team (AUA-600), to explore the feasibility of 
implementing reduced oceanic aircraft separations. These organizations formed an Experimental Working Group to make high- 
level decisions regarding the implementation of the proposed separation standard. In response, ACT-540 formed a Research Team 
to design and conduct a concept exploration study at the FAA William J. Hughes Technical Center. The Research Team led all 
efforts including the planning and design of the simulation and conduct of a simulation. The team also queried the controllers and 
compiled their responses regarding the proposed procedure. This report discusses the Reduced Horizontal Separation Minima 
(RHSM) concept exploration simulation. It describes the simulation, procedures, and tools developed to ascertain the experiences 
of individuals who participated. The concept exploration examined issues that might affect a controller's ability to manage 
reduced longitudinal separation in the oceanic environment. A demonstration of the RHSM concept was conducted in the Oceanic 
Laboratory at the Federal Aviation Administration (FAA) William J. Hughes Technical Center on November 6 and 7, 1996. 
DTIC 
Air Traffic; Controllers; Pacific Ocean 

19980227311 Federal Aviation Administration, Civil Aeromedical Inst., Oklahoma City, OK USA 

The Relationship of Sector Characteristics to Operational Errors Final Report 

Rodgers, Mark D.; Mogford, Richard H.; Mogford, Leslye S.; May 1998; 66p; In English; Prepared in collaboration with William 

J. Hughes Technical Center, Atlantic City, NJ and Rigel Associates, Marmora, NJ. 

Contract(s)/Grant(s) : DTFA02-95-P-35434 

Report No.(s): AD-A350717; DOT/FAA/AM-98/14; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

An exploratory study was conducted on the relationship of air traffic control (ATC) complexity factors to operational errors 
(OEs). This consisted of a detailed examination of OE data from 1992 through 1995 from the Atlanta en route center. The System- 
atic Air Traffic Operations Research Initiative (S ATORI) system was used to collect data for the analysis. Sectors were categorized 
into zero-, low-, and high-error groups. Fifteen sector and traffic flow variables had statistically significant correlations with OE 
frequency. Four variables were higher for the high-error group as compared to the zero-error group. Sector size was smaller for 
the high-error group as compared to the combined zero- and low-error categories. A significant multiple correlation was found 
between overall OE rate and a subset of the ATC complexity measures. The data were also analyzed to define relationships 
between the complexity measures and controller situational awareness (SA) at the time of the OE. The only statistically significant 
difference between OEs with and without SA was for horizontal separation. In addition, high-error sectors were characterized by 
low SA for errors. Certain sector and traffic flow characteristics were associated with these high-error sectors, suggesting that 
these factors may negatively affect SA. It was concluded that the results demonstrated a relationship between sector complexity 
and OE rate. Such findings, if extended, could assist with traffic management, sector design activities, and the development of 
decision-support systems. 
DTIC 
Operations Research; Errors; Air Traffic Control; Error Analysis 

19980227320 Oklahoma Univ., Dept. of Psychology, Norman, OK USA 



Gronlund, Scott D.; Ohrt, Daryl D.; Dougherty, Michael R.; Perry, Jennifer L.; Manning, Carol A.; May 1998; 14p; In English 
Contract(s)/Grant(s) : DTFA02-93-D-93088 

22 



Report No.(s): AD-A350417; DOT/FAA/AM-98/16; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

We tested en route air traffic controllers (currently serving as instructors at the FAA Academy) to determine what they remem- 
ber about the aircraft in their sector. We focused on memory for flight data (especially aircraft altitude and ground speed) and the 
position of the aircraft on the radar screen. Aircraft importance affected memory for flight data but not the highly accurate recall 
of the radar position of the aircraft. We hypothesize that controllers use their excellent memory for aircraft position to classify 
aircraft as important (potential traffic) or not, and better remember flight data about important aircraft (in particular, their exact 
altitude). The results have implications for improving techniques to assess situation awareness and interfaces to support it. 
DTIC 
Air Traffic Controllers (Personnel); Alertness; Memory; Air Traffic Control 

19981)227343 Civil Aeromedical Inst., Oklahoma City, OK USA 

The Combination of Flight Count and Control Time as a New Metric of Air Traffic Control Activity Final Report 

Mills, Scott H.; May 1998; 15p; In English 

Report No.(s): AD-A350504; DOT/FAA/AM-98/15; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The exploration of measures of airspace activity is useful in a number of significant ways, including the establishment of 
baseline air traffic control (ATC) measures and the development of tools and procedures for airspace management. This report 
introduces a new metric of ATC activity that combines two existing measures (flight count and the time aircraft are under control). 
The Aircraft Activity Index (AAI) is sensitive to changes in both flight count and flight length, and therefore is a superior measure 
for comparing aircraft activity between two epochs of time. The AAI was applied to data from 10 days of System Analysis Record- 
ings obtained from the Seattle Air Route Control Center. The advantages of the AAI were most apparent when different aircraft 
types consistently had different mean flight lengths. Possible uses of the AAI and other ATC measures for the evaluation of new 
systems and procedures are discussed. 
DTIC 
Air Traffic Control; Systems Analysis; Airspace 

19980227346 Mississippi State Univ., Aerospace Engineering, Mississippi State, MS USA 

Flight Test Evaluation of a Differential Global Positioning System Sensor io Munway Performance Testing 

Germann, Kenneth Paul; Aug. 04, 1998; 78p; In English 

Report No.(s): AD-A350715; 98-035; No Copyright; Avail: CASI; A05, Hardcopy; A01, Microfiche 

This study discusses the use of a carrier phase differential global positioning system (DGPS) receiver set in basic takeoff and 
landing performance flight testing. A technique for using DGPS receivers as theodolites in takeoff and landing performance tests 
is developed. Both position and velocity data are available from a DGPS receiver. As a result distances can be calculated by differ- 
encing the position coordinates or by integrating the available ground velocities. Both of these techniques are used and compared 
to a traditional video theodolite system for ground roll distances. . The viability of using DGPS ground speed data in lieu of air 
data in calculating the distance to clear a barrier is also explored. These methods are used to determine the nominal takeoff and 
landing performance of an experimental general aviation airplane. Test results are mixed. DGPS velocity integration yields good 
results for ground phase calculations. All other results are inconclusive. 
DTIC 

Flight Tests; Global Positioning System,; Runways; Performance Tests; Receivers; Takeoff; Aircraft Landing; Aircraft 
Performance 

19980227424 Civil Aeromedical Inst., Oklahoma City, OK USA 

An Analysis of Voice Communication io a Simulated Approach Control Environment Final Report 

Prinzo, O. V., Civil Aeromedical Inst., USA; May 1998; 30p; In English 

Report No.(s): AD-A350523; DOT/FAA/AM-98/17; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

This report consists of an analysis of simulated terminal radar approach control (TRACON) air traffic control communica- 
tions. Twenty-four full performance level air traffic controllers (FPLATC) from 2 TRACON facilities participated in the simula- 
tion study. Each controller worked 2 light- and 2 heavy-traffic density scenarios for feeder and final sectors. All communications 
were audio recorded and transcribed verbatim by a retired FPLATC. Once transcribed, transmissions were parsed into commu- 
nication elements. Each communication element was assigned a speech act category (e.g., address, instruction, request, or advi- 
sory), an aviation topic (e.g., altitude, heading, speed) and then coded for irregularities (e.g., grouping numbers together when 
they should be spoken sequentially, or omitting, substituting, or adding words contrary to required phraseology) (ATSAT, Prinzo 
et al., 1995). The simulated communications were compared to an analysis performed on audiotapes from the same TRACON 
facilities. Percentages in 3 speech act categories were comparable (Instruction, 55% versus 51%; Address; 14% versus 26%; Advi- 

23 



sory, 24% versus 18%). Detailed analyses revealed that, although there were fewer irregular communications produced during 

simulation, the distributions of those communication irregularities were very much the same, with the exception of aircraft call 

sign. The differences in those distributions were attributed to the voice recognition system; it could not recognize a call sign spoken 

sequentially and then restated in grouped form. 

DTIC 

Voice Communication; Simulation; Controllers; Radar Approach Control; Air Traffic Controllers (Personnel); Air Traffic 

Control 



05 
AIRCRAFT DESIGN, TESTING AND PERFORMANCE 

Includes aircraft simulation technology. 

19980221786 Federal Aviation Administration, Fire Safety Section, Atlantic City, NJ USA 

Cargo Compartment Fire Protection in Large Commercial Transport Aircraft 

Blake, D., Federal Aviation Administration, USA; Marker, T., Federal Aviation Administration, USA; Hill, R., Federal Aviation 

Administration, USA; Reinhardt, J., Federal Aviation Administration, USA; Sarkos, C, Federal Aviation Administration, USA; 

Jul. 1998; 30p; In English 

Report No.(s): PB98-163298; DOT/FAA/AR-TN98/32; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

This report describes recent research by the Federal Aviation Administration (FAA) related to cargo compartment fire protec- 
tion in large transport aircraft. A gaseous hydrofluorocarbon, HFC-125, was compared to Halon 1301 in terms of fire suppression 
effectiveness and agent decomposition levels in the cargo compartment and passenger cabin during full-scale tests involving a 
bulk-loaded cargo fire. Also, a zoned water mist system was designed and evaluated against a bulk-loaded cargo fire. An exploding 
aerosol can simulator is being developed to provide a repeatable fire threat for evaluation of new halon replacements agents. The 
potential severity of an exploding aerosol can inside a cargo compartment and the effectiveness of Halon 1 301 inerting was dem- 
onstrated. Tests were also conducted to determine the effectiveness of Halon 1201 against a carbon fire involving oxygen canisters. 
Finally, HFC-125 was evaluated for use as a simulant for Halon 1301 during cargo compartment approval testing to demonstrate 
compliance with applicable FAA regulations. 
NTIS 
Commercial Aircraft; Fire Prevention; Transport Aircraft; Aerosols 



19981)221795 Federal Aviation Administration, Fire Safety Section, Atlantic City, NJ USA 

Effects of Concentrated Hydrochloric Acid Spills stnd Aircraft Aluminum Skin 

Speitel, L. C, Federal Aviation Administration, USA; Jul. 1998; 18p; In English 

Report No.(s): PB98-163280; DOT/FAA/AR-TN97/108; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The purpose of this study is to evaluate the effect of a spill of concentrated hydrochloric acid (HCL) on the aircraft aluminum 
skin of a cargo compartment and to determine the time required for a spill to cause catastrophic failure for a worst-case scenario. 
NTIS 
Hydrochloric Acid; Aluminum,; Cargo 



19980223583 NASA Langley Research Center, Hampton, VA USA 

Some Information of the Operational Experiences of Turbine-Powered Commercial Transports 

Jewel, Joseph W., Jr., NASA Langley Research Center, USA; Hunter, Paul A., NASA Langley Research Center, USA; McLaugh- 
lin, Milton D., NASA Langley Research Center, USA; Jul. 20, 1961; 26p; In English 
Report No.(s): NASA-TM-SX-595; L-1696; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

This report presents a brief discussion of some information on the operational experiences noted on VGH records from six 
types of turbine- powered commercial transport aircraft. These flight characteristics cover oscillatory motions, maneuver accel- 
erations, sinking speeds, placard speed exceedances, and miscellaneous or unusual flight events. 
Author 
Commercial Aircraft; Flight Characteristics; Turbine Engines 

24 



19980223698 NASA Lewis Research Center, Cleveland, OH USA 

Analytical Steely of Soft Landings on Gas-Filled Bags 

Esgar, Jack B., NASA Lewis Research Center, USA; Morgan, William C, NASA Lewis Research Center, USA; Jan. 01, 1960; 

32p; In English 

Report No.(s): NASA-TR-R-75; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An analytical procedure was developed that is valid for bags of various arbitrary shapes and is applicable to planetary or lunar 
landings for sinking speeds that are small compared to the sonic velocity of the gas within the bag. For landing on the earth at speeds 
consistent with normal parachute descent, the relative merits of four bag shapes were evaluated both with and without gas bleed 
from the bags. Deceleration and onset rates acceptable for well-supported humans seem feasible. 
Author (revised) 
Soft Landing; Gas Bags; Parachute Descent; Descent Trajectories 

1998022361.2 NASA Langley Research Center, Hampton, VA USA 

Status of Spin Research for Recent Airplane Designs 

Neihouse, Anshal I., NASA Langley Research Center, USA; Klinar, Walter J., NASA Langley Research Center, USA; Scher, 

Stanley H., NASA Langley Research Center, USA; 1960; 58p; In English 

Report No.(s): NASA-TR-R-57; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

This report presents the status of spin research for recent airplane designs as interpreted at the Langley Research Center of 
the National Aeronautics and Space Administration. Major problem areas discussed include: (1) Interpretation of results of spin- 
model research (2) Analytical spin studies (3) Techniques involved in obtaining measurements of various parameters in the spin 
(4) Effectiveness of controls during spins and recoveries (5) Influence of long noses, strakes, and canards on spin and recovery 
characteristics (6) Correlation of spin and recovery characteristics for recent airplane and model designs. Analyses conclusions 
are drawn. 
Author 
Aircraft Spin; Spin Dynamics; Aerodynamic Characteristics; Aerodynamic Stalling; Aircraft Stability; Flight Characteristics 

19980223919 NASA Dryden Flight Research Center, Edwards, CA USA 
An Overview of an Experimental Demonstration Aerotow Program 

Murray, James E., NASA Dryden Flight Research Center, USA; Bowers, Albion H., NASA Dryden Flight Research Center, USA; 
Lokos, William A., NASA Dryden Flight Research Center, USA; Peters, Todd L., NASA Dryden Flight Research Center, USA; 
Gera, Joseph, Analytical Services and Materials, Inc., USA; Sep. 1998; 29p; In English; 30th, 15-17 Sep. 1998, Reno, NV, USA; 
Sponsored by Society of Flight Test Engineers, USA 
Contact(s)/Grant(s): RTOP 242-33-02-00-25 

Report No.(s): NASA/TM- 1998-206566; H-2279; NAS 1.15:206566; No Copyright; Avail: CASI; A03, Hardcopy; A01, Micro- 
fiche 

An overview of an experimental demonstration of aerotowing a delta- wing airplane with low -aspect ratio and relatively high 
wing loading is presented. Aerotowing of future space launch configurations is a new concept, and the objective of the work 
described herein is to demonstrate the aerotow operation using an airplane configuration similar to conceptual space launch 
vehicles. Background information on the use of aerotow for a space launch vehicle is presented, and the aerotow system used in 
this demonstration is described. The ground tests, analytical studies, and flight planning used to predict system behavior and to 
enhance flight safety are detailed. The instrumentation suite and flight test maneuvers flown are discussed, preliminary perfor- 
mance is assessed, and flight test results are compared with the preflight predictions. 
Author 
Tetherlines; Towing; Tethering; Delta Wings; Flight Tests; Launch Vehicles; Low Aspect Ratio; Wing Loading 

19980223930 Naval Air Warfare Center, Aircraft Div., Patuxent River, MD USA 

Flight Test Automation Options 

Carico, Dean, Naval Air Warfare Center, USA; 1998; 13p; In English 

Report No.(s): AD-A350677; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Flight testing is often called the key component of test and evaluation. The cost of conventional flight testing is expected to 
escalate approaching the 21st century and beyond. Augustine noted several years ago that if this trend continues, a single advanced 
fighter aircraft would cost more than the entire DoD budget by the middle of next century. As the cost of flight testing continues 
to escalate in a predicted hostile fiscal environment, it is important to consider options to help minimize flight test cost. Sugges- 
tions range from completely eliminating developmental testing to employing a variety of flight test automation options. Flight 

25 



test automation option concepts range from the fantasy of "push a button, the test is done," to the more practical use of a personal 
computer to help with some repetitive flight test tasks and to help store large amounts of related data. Options to help automate 
specific aspects of flight testing are starting to gain acceptance. Several test automation options exist that have the potential to 
enhance flight testing by permitting it to be done better, faster, cheaper, and safer. This paper briefly discusses a variety of flight 
test automation options including the OSD Automated Test Planning System (ATPS) work to automate the test and evaluation 
master plan (TEMP), the Army Test and Evaluation Planning and Reporting System (TEPRS), the G&C System work on Test 
DTIC 
Flight Tests; Data Acquisition; Test Ranges; Data Storage 

19980223931 Naval Air Warfare Center, Aircraft Div., Patuxent River, MD USA 
Telemetry and HPC Potential Applications 

Normyle, Dennis, Naval Air Warfare Center, USA; 1998; lOp; In English 

Report No.(s): AD-A350675; No Copyright; Avail: CASI; A02, Hardcopy; A01, Microfiche 

The intent of this paper is to familiarize the reader with the telemetry world and to investigate possible applications that the 
High Performance Computer (HPC) facility will provide for the T&E community. The following is a brief outline of the paper. 
1) Current NAWC-AD Telemetry Capability - This section describes the current capability of RTPS (i.e. Number of PES rooms, 
Room Layout, Computers Used, Projects Supported, Number of flights flown etc.... ). 2) Anatomy of a F/A-18E1 Flutter flight 
- This section will describe the anatomy of a flutter flight, the type of maneuvers performed, the type of data collected, the 
AMOUNT of data collected, and how the data is processed post flight. 3) Telemetry/A CETEF Applications. A brief discussion 
on how RTPS and Manned Flight Simulator are currently linked and some of the early applications used. 4) Possible applications 
for HPC in Telemetry applications. 5) Conclusion. 
DTIC 
Telemetry; Utilization; Flight Tests 

19980223932 Naval Air Warfare Center, Aircraft Div., Patuxent River, MD USA 
Future Naval UCAV Applications & Enabling Technologies 

Booz, Julieta E., Naval Air Warfare Center, USA; 1998; 12p; In English 

Report No.(s): AD-A350673; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Briefing notes on the applications and enabling technologies for future naval Unmanned Combat Air Vehicles. 
DTIC 
Remotely Piloted Vehicles; Technologies 

19980223943 Naval Air Warfare Center, Aircraft Div., Patuxent River, MD USA 
Naval Rotary Wing Aircraft Flight Test Squadron Flight Test Approval Process 
Mertaugh, Lawrence J., Naval Air Warfare Center, USA; Jan. 1998; 6p; In English 
Report No.(s): AD-A350674; No Copyright; Avail: CASI; A02, Hardcopy; A01, Microfiche 

This presentation will provide a description of the process used by the Naval Rotary Wing Aircraft Test Squadron, at Patuxent 
River, for minimizing the risk associated with its flight test operations. This process is defined in terms of three basic functions. 
These functions are: Test Plan, Flight Clearance, and the Aircraft Modification/Configuration Control Sheet. It is through these 
functions that we provide oversight of the test planning, insure that any required aircraft modifications are sound, and provide 
controls over the aircraft modification process. Each of these functions play a role throughout the test program in preventing 
changes in testing that could jeopardize the quality of the test results or the safety of the crew or the aircraft. 
DTIC 
Rotary Wing Aircraft; Flight Tests; Risk 

19980223965 NASA Langley Research Center, Hampton, VA USA 

In-Flight System Identification 

Morelli, Eugene A., NASA Langley Research Center, USA; 1998; lOp; In English; Atmospheric Flight Mechanics, 10-12 Aug. 

1998, Boston, MA, USA; Sponsored by American Inst, of Aeronautics and Astronautics, USA 

Report No.(s): AIAA Paper 98-4261; No Copyright; Avail: Issuing Activity, Hardcopy, Microfiche 

A method is proposed and studied whereby the system identification cycle consisting of experiment design and data analysis 
can be repeatedly implemented aboard a test aircraft in real time. This adaptive in-flight system identification scheme has many 
advantages, including increased flight test efficiency, adaptability to dynamic characteristics that are imperfectly known a priori, 
in-flight improvement of data quality through iterative input design, and immediate feedback of the quality of flight test results. 

26 



The technique uses equation error in the frequency domain with a recursive Fourier transform for the real time data analysis, and 

simple design methods employing square wave input forms to design the test inputs in flight. Simulation examples are used to 

demonstrate that the technique produces increasingly accurate model parameter estimates resulting from sequentially designed 

and implemented flight test maneuvers. The method has reasonable computational requirements, and could be implemented 

aboard an aircraft in real time. 

Author 

Design Analysis; Flight Tests; Dynamic Characteristics; Feedback 



19980223968 NASA Langley Research Center, Hampton, VA USA 

Subsonic Flight Teste of a 1/7-Scale Radio-Controlled Model of the North American X-15 Airplane with Particular Refer- 
ence to High Angel-of-Attack Conditions 

Hewes, Donald E., NASA Langley Research Center, USA; Hassell, James L., Jr., NASA Langley Research Center, USA; Jun. 
1960; 46p; In English 
Report No.(s): NASA-TM-X-283; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An investigation of the subsonic stability and control characteristics of an unpowered 1/7-scale model based on the North 
American X-15 airplane was conducted by using a radio-controlled model launched from a helicopter and flown in free-gliding 
flight. At angles of attack below about 20 deg. where the model motions represent those of the X-15 airplane, the model was found 
to be both longitudinally and laterally stable, and the all-movable tail surfaces were found to be very effective. The model could 
also be flown at much higher angles of attack where the model motions did not necessarily represent those of the airplane because 
of slight geometrical differences and Reynolds number effects, but these test results are useful in evaluating the effectiveness at 
these angles of the type of lateral control system used in the X-15 airplane. In some cases, the model was flown to angles of attack 
as high as 60 or 70 deg. without encountering divergent or uncontrollable conditions. For some flights in which the model was 
subjected to rapid maneuvers, spinning motions were generated by application of corrective controls to oppose the direction of 
rotation. Rapid recoveries from this type of motion were achieved by applying roll control in the direction of rotation. 
Author 
Subsonic Speed; Flight Tests; Scale Models; Angle of Attack; Free Flight; Aerodynamic Stability; Aircraft Control; X-15 Aircraft 



19980223972 NASA Langley Research Center, Hampton, VA USA 

Aerodynamic Characteristics of a Target: Drone Vehicle at Macfa Numbers from 1,57 to 2,10 

Blair, A. B., Jr., NASA Langley Research Center, USA; Founder, Roger H., NASA Langley Research Center, USA; Jul. 1968; 

72p; In English 

Report No.(s): NASA-TM-SX-1531; AF-AM-627; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

An investigation of a 1/4-scale supersonic target drone model was performed in the Langley Unitary Plan wind tunnel to deter- 
mine the effects of various sizes of canards, vertical tails, and ailerons on the aerodynamic characteristics. The tests were made 
at Mach numbers from 1.57 to 2. 10 through an angle-of -attack range from about -5 to 23 deg. 
Author 

Aerodynamic Characteristics; Aircraft Structures; Control Surfaces; Supersonic Speed; Wind Tunnel Tests; Scale Models; Drone 
Vehicles; Wind Tunnel Models 



r 3 NASA Langley Research Center, Hampton, VA USA 

Characteristics of a Revised Target Drone Vehicle at Mach Numbers from 1.60 to 2.86 
Blair, A. B., Jr., NASA Langley Research Center, USA; Babb, C. Donald, NASA Langley Research Center, USA; Feb. 1968; 54p; 
In English 
Report No.(s): NASA-TM-SX-1532; L-5824; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

An investigation has been conducted in the Langley Unitary Plan wind tunnel to determine the aerodynamic characteristics 
of a revised target drone vehicle through a Mach number range from 1.60 to 2.86. The vehicle had canard surfaces and a swept 
clipped-delta wing with twin tip-mounted vertical tails. 
Author 

Aerodynamic Characteristics; Delta Wings; Canard Configurations; Drone Vehicles; Swept Wings; Wind Tunnel Tests; Super- 
sonic Speed; Wind Tunnel Models; Aerodynamic Stability 

27 



19980223993 NASA Ames Research Center, Moffett Field, CA USA 

Flight Investigation of the Low-Speed Characteristics of a 45 deg Swept- Wing Fighter-Type Airplane with Blowing 

Boundary-Layer Control Applied to the Leading- and Trailing-Edge Flaps 

Quigley, Hervey C, NASA Ames Research Center, USA; Anderson, Seth B., NASA Ames Research Center, USA; Innis, Robert 

C, NASA Ames Research Center, USA; Sep. 1960; 46p; In English 

Report No.(s): NASA-TN-D-321; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A flight investigation has been conducted to study how pilots use the high lift available with blowing-type boundary-layer 
control applied to the leading- and trailing-edge flaps of a 45 deg. swept- wing airplane. The study includes documentation of the 
low-speed handling qualities as well as the pilots' evaluations of the landing-approach characteristics. All the pilots who flew the 
airplane considered it more comfortable to fly at low speeds than any other F-100 configuration they had flown. The major 
improvements noted were the reduced stall speed, the improved longitudinal stability at high lift, and the reduction in low-speed 
buffet. The study has shown the minimum comfortable landing-approach speeds are between 120.5 and 126.5 knots compared 
to 134 for the airplane with a slatted leading edge and the same trailing-edge flap. The limiting factors in the pilots' choices of 
landing-approach speeds were the limits of ability to control flight-path angle, lack of visibility, trim change with thrust, low static 
directional stability, and sluggish longitudinal control. Several of these factors were found to be associated with the high angles 
of attack, between 13 deg. and 15 deg., required for the low approach speeds. The angle of attack for maximum lift coefficient 
was 28 deg. 
Author 

Swept Wings; Fighter Aircraft; Boundary Layer Control; Externally Blown Flaps; Aerodynamic Coefficients; Aerodynamic Char- 
acteristics; Longitudinal Stability; Directional Stability 

19980227097 NASA Flight Research Center, Edwards, CA USA 

Flight Investigation of the Lift and Drag Characteristics of a Swept-Wing, Multijet, Transport-Type Airplane 

Tambor, Ronald, NASA Flight Research Center, USA; Sep. 1960; 28p; In English 

Report No.(s): NASA-TN-D-30; H-119; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The lift and drag characteristics of a Boeing KC-135 airplane were determined during maneuvering flight over the Mach num- 
ber range from 0.70 to 0.85 for the airplane in the clean configuration at an altitude of 26,000 feet. Data were also obtained over 
the speed range of 130 knots to 160 knots at 9,000 feet for various flap deflections with gear down. 
Author 
C-135 Aircraft; Swept Wings; Lift; Aerodynamic Drag; Subsonic Speed; Aerodynamic Characteristics; Jet Aircraft; Flight Tests 

19980227163 Army Command and General Staff Coll., Fort Leavenworth, KS USA 
The Implications of Video Dataiink on the AC-130, 5 Aug. 1997 - 5 Jun. 1998 

Hicks, John M., Army Command and General Staff Coll., USA; Jun. 05, 1998; 95p; In English 
Report No.(s): AD-A350132; No Copyright; Avail: CASI; A05, Hardcopy; A01, Microfiche 

This study considers the implications of video dataiink (VDL) on the AC-130. Gunships use infrared and low-light television 
sensors, and synthetic aperture radar to search for and to identify target for close air support and interdiction missions. The addition 
of VDL offers gunship crews the ability to employ real-time information to the cockpit/offboard targeting (RTIC/OT) technology 
to improve situational awareness, survivability, and operational flexibility. Also, VDL offers the joint force air component com- 
mander (JFACC) inflight tasking capability, increased reconnaissance capability, operational flexibility and situation awareness. 
Ultimately, VDL allows command and control elements to exercise direct control of gunship operations. These capabilities are 
beneficial when they provide information to the crew or to the JFACC. However, VDL used to provide direct control of gunship 
operations may violate the Air Force doctrinal tenet of centralized control and decentralized execution. Lessons learned from 
recent contingencies, leadership doctrine, academic works on leadership and management theory all suggest that direct control 
of tactical mission can cause decreased survivability, ineffective span of control, task saturation, tactical inflexibility, mistrust 
between commanders and subordinates, decreased morale, and subordinates that lack initiative. The study provides recommenda- 
tions to mitigate potential problems associated with the use of VDL on gunships. 
DTIC 
Data Links; Video Data; Television Systems; Alternating Current; Information Transfer; Infrared Detectors ; Video Signals 

19980227164 Army Command and General Staff Coll., Fort Leavenworth, KS USA 
An Analysis of Prime Vendor Support for the AII64 Apache 

Angelo, Anthony W., Army Command and General Staff Coll., USA; Jun. 05, 1998; 95p; In English 
Report No.(s): AD-A350090; No Copyright; Avail: CASI; A05, Hardcopy; A01, Microfiche 

28 



This study investigates the use of prime vendor support for the Army's AH64 Apache helicopter. It defines the term prime 
vendor support and it analyzes the reasons for applying this concept of logistical support to the Army's aviation support doctrine. 
This study shows why privatization of supply parts management has become not only necessary, but a driving force in the develop- 
ment and future application of prime vendor support. This study concludes that prime vendor support can create an innovative 
partnership between the Army and the Apache's prime vendor that will minimize the time it takes to deliver parts to mechanics 
and delay the purchasing of parts until they are needed to complete repairs. However, as the Army pursues a strategy of transforma- 
tion needed to get from today's multiecheloned logistics system to more streamlined and efficient processes of support it must 
proceed with extreme caution. The complexities of resource management and the effects of changing existing processes to new 
concepts of support mandate further analysis and the development of procedures that will mitigate the risks associated with prime 
vendor support. 
DTIC 
Helicopters; AH-64 Helicopter; Investigation; Logistics Management 

19980227171 Army Command and General Staff Coll., Fort Leavenworth, KS USA 

Air Superiority Fighter Characteristics 

Browne, James S., Army Command and General Staff Coll., USA; Jun. 05, 1998; 106p; In English 

Report No.(s): AD-A350022; No Copyright; Avail: CASI; A06, Hardcopy; A02, Microfiche 

This study determines the essential characteristics of an air superiority fighter. Its importance stems from the assumption that 
air superiority is paramount in any military operation and that fighter aircraft play a major role. Air superiority as well as roles, 
functions, and missions are defined in chapter one to develop an understanding of the operative terms and definitions used through- 
out the thesis. This thesis is an in-depth study of the historical characteristics of the air superiority fighter. A complete review of 
air superiority fighter evolution is divided into four distinct generations. The review includes example aircraft that highlight the 
consistent characteristics found in each generation. The thesis research and analysis chapters focus on three key areas of interest. 
They are: (1) aircraft design, (2) avionics and weapons, and (3) training. The key areas of interest are coupled with a discussion 
of cost considerations during analysis. Fiscal constraints are a major factor in design and employment limitations. The thesis con- 
cludes that there are three essential characteristics of an air superiority fighter: (1) the aircraft is designed for the air-to-air role, 
(2) the aircraft has the first launch opportunity, and (3) the aircraft is flown by singularly trained air-to-air pilots. 
DTIC 
Cost Analysis; Aircraft Design; Fighter Aircraft 

19980227179 NASA Langley Research Center, Hampton, VA USA 

Summary of V-G and VGH Data Collected on Lockheed Electra Airplanes During Airplane Operations 

Jewel, Joseph W., Jr., NASA Langley Research Center, USA; Fetner, Mary W., NASA Langley Research Center, USA; 1961; 60p; 

In English 

Report No.(s): NASA-TM-SX-523; L-1467; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

Data obtained by NASA VGH and V-G recorders on several Lockheed Electra airplanes operated over three domestic routes 
have been analyzed to determine the in-flight accelerations, airspeed practices, and landing accelerations experienced by this par- 
ticular airplane. The results indicate that the accelerations caused by gusts and maneuvers are comparable to corresponding results 
for piston-engine transport airplanes. Oscillatory accelerations (apparently caused by the autopilot or control system) appear to 
occur about one-tenth as frequently as accelerations due to gusts. Airspeed operating practices in rough air generally follow the 
trends shown by piston-engine transports in that there is no significant difference between the average airspeed in rough or smooth 
air. Placard speeds were exceeded more frequently by the Electra airplane than by piston-engine transport airplanes. Generally, 
the landing-impact accelerations were higher than those for piston-engine transports. 
Author 
Transport Aircraft; Commercial Aircraft; Airline Operations; Impact Loads; Landing Loads 

19980227196 NASA Langley Research Center, Hampton, VA USA 

Summary of Flight-Test Results of the VZ-2 Tilt- Wing Aircraft 

Pegg, Robert J., NASA Langley Research Center, USA; Feb. 1962; 44p; In English 

Report No.(s): NASA-TN-D-989; L-1574; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Flight-test information gained from a tilt-wing research aircraft tested at the Langley Research Center has shown that design 
problems exist in such fields as low-speed stability and control, handling qualities, and flow separation during transition. The con- 
trol power in the near-hovering configuration was considered by the pilots to be inadequate in yaw, marginal in pitch, and excessive 
in roll. Solutions for some of the design problems are indicated; for example, the addition of a leading-edge droop to the wing 

29 



in an attempt to delay flow separation resulted in such significantly improved handling qualities in the transition range that an 

additional descent capability of 1,100 feet per minute was obtained. 

Author 

Flight Tests; Boundary Layer Separation; Vz-2 Aircraft; Separated Flow; Design Analysis; Leading Edges; Hovering 

19980227205 NASA Ames Research Center, Moffett Field, CA USA 

Aerodynamic Performance and Static Stability at Mach Number 3.3 of an Aircraft Configuration Employing Three Trian- 
gular Wing Panels and a Body Equal Length 

James, Carlton S., NASA Ames Research Center, USA; Aug. 1960; 40p; In English 
Report No.(s): NASA-TN-D-330; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An aircraft configuration, previously conceived as a means to achieve favorable aerodynamic stability characteristics., high 
lift-drag ratio, and low heating rates at high supersonic speeds., was modified in an attempt to increase further the lift-drag ratio 
without adversely affecting the other desirable characteristics. The original configuration consisted of three identical triangular 
wing panels symmetrically disposed about an ogive-cylinder body equal in length to the root chord of the panels. This configura- 
tion was modified by altering the angular disposition of the wing panels, by reducing the area of the panel forming the vertical 
fin, and by reshaping the body to produce interference lift. Six-component force and moment tests of the modified configuration 
at combined angles of attack and sideslip were made at a Mach number of 3.3 and a Reynolds number of 5.46 million. A maximum 
lift-drag ratio of 6.65 (excluding base drag) was measured at a lift coefficient of 0.100 and an angle of attack of 3.60. The lift-drag 
ratio remained greater than 3 up to lift coefficient of 0.35. Performance estimates, which predicted a maximum lift-drag ratio for 
the modified configuration 27 percent greater than that of the original configuration, agreed well with experiment. The modified 
configuration exhibited favorable static stability characteristics within the test range. Longitudinal and directional centers of pres- 
sure were slightly aft of the respective centroids of projected plan-form and side area. 
Author 

Aerodynamic Characteristics; Aircraft Configurations; Wing Panels; Aerodynamic Stability; Lift Drag Ratio; Delta Wings; Aero- 
dynamic Drag; Aerodynamic Coefficients 

19980227208 Naval Aerospace Medical Research Lab., Pensacola, FL USA 

The Development and Initial Validation of the Unmanned Aerial Vehicle (UAV) External Pilot Selection System 

Biggerstaff, S., Naval Aerospace Medical Research Lab., USA; Blower, D. J., Naval Aerospace Medical Research Lab., USA; 

Portman, C. A., Naval Aerospace Medical Research Lab., USA; Chapman, A. D., Naval Aerospace Medical Research Lab., USA; 

Mar. 05, 1998; 21p; In English 

Report No.(s): AD-A350547; NAMRL-1398; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The purpose of this study was to develop physical and selection performance standards for the screening of candidates for 
entrance into the Unmanned Aerial Vehicle (UAV) Pioneer Pilot training program. A minimum Pioneer crew consists of an exter- 
nal pilot, internal pilot, and a mission commander/payload specialist. The mission commander/payload specialist is responsible 
for the overall planning and execution of the specific mission and control of the visual/information gathering during the mission. 
The internal pilot is responsible for the control of the Pioneer when it is beyond visual range. The external pilot is responsible for 
takeoffs, landings, and any in- visual-range control of the vehicle. A task analysis was done in the training and fleet squadrons 
to identify critical tasks for safe flight and the relevant skills required to perform the piloting tasks. From this task analysis, specific 
computer-based tests batteries were chosen as potential predictor variables. The system was programmed and students and exter- 
nal pilots were administered the test battery. A composite training measure was created from objective training scores, verified 
with subjective instructor ratings, and used as the criterion for predictive validation of the system. The sample size was small for 
the preliminary model, but a significant relationship between a composite of multitask tracking scores and UAV performance was 
observed (adjusted R2 = 0.86). In addition, structured and unstructured interviews of the Pioneer crews, students, instructors and 
senior squadron personnel were used to identity important physical characteristics essential for safe operation of the Pioneer. 
These traits were then used to derive medical screening criteria for all crew positions. 
DTIC 
Pilot Selection; Pilot Performance; Pilot Training; Proving; Pilotless Aircraft 

19980227272 Naval Aerospace Medical Research Lab., Pensacola, FL USA 

Landing Craft Air Cushion (LCAC) Navigator Selection System; Initial Model Development 

Biggerstaff, S., Naval Aerospace Medical Research Lab., USA; Blower, D. J., Naval Aerospace Medical Research Lab., USA; 

Portman, C. A., Naval Aerospace Medical Research Lab., USA; Chapman, A., Naval Aerospace Medical Research Lab., USA; 

Mar. 05, 1998; 27p; In English 

30 



Report No.(s): AD-A350546; NAMRL-1399; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The LCAC is an amphibious hovercraft that can ride on a cushion of air across land or sea. Its control features are similar 
to a helicopter and it is designed to transport weapons, cargo, equipment and combat personnel. In the 1980s, the LCAC commu- 
nity was experiencing a high attrition rate, partially due to the absence of any valid selection mechanisms for crewmembers. The 
Naval Aerospace Medical Research Laboratory (NAMRL) developed a selection system for the LCAC operators and engineers 
and the system was transitioned to the Naval Operational Medicine Institute (NOMI) in 1992. Similar attrition problems were seen 
in the Navigator community and in FY 94-95 NAMRL was again tasked with developing a selection system. Concurrent validation 
of the system was done using 58 LCAC navigators. The preliminary predictive model was generated and the cut-off score derived 
from the sponsor's operational manpower needs, to date, 30 candidates have been screened with 25 being recommended for train- 
ing. Thirteen of these candidates have entered training and completed the lull 22 week syllabus. Five of the thirteen recommended 
candidates attrited during training, with multi -tasking as the main reason sited. The system is still being evaluated to: (1) possibly 
include more multi-tasking tests, (2) modify the predictive algorithm, and/or (3) raise the cut-off score to further reduce the attri- 
tion rates. 
DTIC 
Air Cushion Landing Systems; Navigators; Cushions; Ground Effect Machines 



19980227289 NASA Langley Research Center, Hampton, VA USA 

Effects of Boattailing and Nozzle Extension on the Thrust-Minus-Drag of a Multiple-Jet Configuration 

Scott, William R., NASA Langley Research Center, USA; Jun. 1961; 50p; In English 

Report No.(s): NASA-TN-D-887; L-862; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A wind-tunnel investigation of the effects of both boattailing and nozzle extension on the thrust-minus-drag of clustered-jet 
configurations has been conducted at Mach numbers from 0.60 to 1 .40 and jet total-pressure ratios from 3 to 20. Three different 
boattails were tested: an 8 deg conical afterbody, a 16 deg circular-arc afterbody, and a third afterbody having a linear area variation 
with length. A cylindrical afterbody also was tested for comparison purposes. Extending from these bodies are four circular jet 
nozzles with a design Mach number of 2.5 which were spaced symmetrically about the body center line. The results indicated that 
an 8 deg conical afterbody provided the highest net thrust efficiency factors of the four models tested when the nozzle exits were 
at the optimum longitudinal location in each case. The other afterbodies in order of decreasing performance were the 16 deg circu- 
lar-arc, the straight-line-area-distribution, and the cylindrical. 
Author 
Aerodynamic Drag; Thrust; Afterbodies; Boattails; Pressure Ratio; Cylindrical Bodies 



19980227282 NASA Dryden Flight Research Center, Edwards, CA USA 

Analysis of X-15 Landing Approach and Flare Characteristics Determined from the First 30 Flights 

Matranga, Gene J., NASA Dryden Flight Research Center, USA; Jul. 1961; 54p; In English 

Report No.(s): NASA-TN-D-1057; H-221; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

The approach and flare maneuvers for the first 30 flights of the X-15 airplane and the various control problems encountered 
are discussed. The results afford a relatively good cross section of landing conditions that might be experienced with future glide 
vehicles having low lift-drag ratios. Flight-derived drag data show that preflight predictions based on wind-tunnel tests were, in 
general, somewhat higher than the values measured in flight. Depending on configuration, the peak lift-drag ratios from flight 
varied from 3.5 to 4.5 as compared with a predicted range of from 3.0 to 4.2. by employing overhead, spiral-type patterns begin- 
ning at altitudes as high as 40,000 feet, the pilots were consistently able to touch down within about +/-1,000 feet of a designated 
point. A typical flare was initiated at a "comfortable" altitude of about 800 feet and an indicated airspeed of approximately 300 
knots., which allowed a margin of excess speed. The flap and gear were extended when the flare was essentially completed, and 
an average touchdown was accomplished at a speed of about 185 knots indicated airspeed, an angle of attack of about 7 deg, and 
a rate of descent of about 4 feet per second. In general, the approach and landing characteristics were predicted with good accuracy 
in extensive preflight simulations. F-104 airplanes which simulated the X-15 landing characteristics were particularly valuable 
for pilot training. 
Author 
X-15 Aircraft; Aerodynamic Characteristics; Angle of Attack; Lift Drag Ratio; Landing; Touchdown; Descent; Approach 

31 



19980227397 NASA Ames Research Center, Moffett Field, CA USA 

Experimental Investigation of a Hypersonic Glider Configuration at si Mach Number of 6 and at Full-Scale Reynolds 

Numbers 

Seiff, Alvin, NASA Ames Research Center, USA; Wilkins, Max E., NASA Ames Research Center, USA; Jan. 1961; 74p; In 

English 

Report No.(s): NASA-TN-D-341; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

The aerodynamic characteristics of a hypersonic glider configuration, consisting of a slender ogive cylinder with three highly 
swept wings, spaced 120 apart, with the wing chord equal to the body length, were investigated experimentally at a Mach number 
of 6 and at Reynolds numbers from 6 to 16 million. The objectives were to evaluate the theoretical procedures which had been 
used to estimate the performance of the glider, and also to evaluate the characteristics of the glider itself. A principal question 
concerned the viscous drag at full-scale Reynolds number, there being a large difference between the total drags for laminar and 
turbulent boundary layers. It was found that the procedures which had been applied for estimating minimum drag, drag due to 
lift, lift curve slope, and center of pressure were generally accurate within 10 percent. An important exception was the non-linear 
contribution to the lift coefficient which had been represented by a Newtonian term. Experimentally, the lift curve was nearly 
linear within the angle-of -attack range up to 10 deg. This error affected the estimated lift-drag ratio. The minimum drag measure- 
ments indicated that substantial amounts of turbulent boundary layer were present on all models tested, over a range of surface 
roughness from 5 microinches maximum to 200 microinches maximum. In fact, the minimum drag coefficients were nearly inde- 
pendent of the surface smoothness and fell between the estimated values for turbulent and laminar boundary layers, but closer 
to the turbulent value. At the highest test Reynolds numbers and at large angles of attack, there was some indication that the skin 
friction of the rough models was being increased by the surface roughness. At full-scale Reynolds number, the maximum lift-drag 
ratio with a leading edge of practical diameter (from the standpoint of leading-edge heating) was 4.0. The configuration was stati- 
cally and dynamically stable in pitch and yaw, and the center of pressure was less than 2-percent length ahead of the centroid of 
plan-form area. 
Author 

Hypersonic Gliders; Aerodynamic Characteristics; Aerodynamic Coefficients; Swept Wings; Hypersonic Speed; Laminar Bound- 
ary Layer; Turbulent Boundary Layer 

19980227332 NASA Langley Research Center, Hampton, VA USA 

f Effects of Flexibility on Wing Strains in Rough Air for a Large Swept- Wing Airplane by Means ofExper- 
)etermined Frequency-Response Functions with an Assessment of Random-Process Techniques Employed 
Coleman, Thomas L., NASA Langley Research Center, USA; Press, Harry, NASA Langley Research Center, USA; Meadows, 
May T, NASA Langley Research Center, USA; 1960; 36p; In English 
Report No.(s): NASA-TR-R-70; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Flight test measurements on a large swept- wing bomber airplane through rough air at altitudes of 5,000 and 35,000 feet are 
analyzed in order to determine the effects of airplane flexibility on wing bending and shear strains. For this purpose, the power 
spectra of the strain responses and the frequency-response functions for the strain responses to vertical gust disturbances are deter- 
mined and compared with the strain responses for a quasi-rigid airplane. The measured power spectra and frequency-response 
functions are subject to distortion and statistical sampling errors from a variety of sources. A general analysis of the reliability 
of such results is presented and methods of estimating the distortions and sampling errors are developed. These methods are 
applied to the interpretation of the test results. 
Author 
Flight Tests; Errors; Sampling; Reliability; Distortion; Frequency Response 

19980227344 Federal Aviation Administration, Airworthiness Assurance Research and Development Branch, Atlantic City, NJ 
USA 

Vertical Drop Test of a Beechcraft 1900C Airliner Final Report, Jul - Nov. 1995 
McGuire, Robert J.; Vu, Tong; May 1998; 83p; In English 

Report No.(s): AD-A350509; AAR-431; DOT/FAA/AR-96/119; No Copyright; Avail: CASI; A05, Hardcopy; A01, Microfiche 
A commuter category Beechcraft 1900C airliner was subjected to a vertical impact drop test at the FA A William J. Hughes 
Technical Center, Atlantic City International Airport, New Jersey. The purpose of this test was to measure the impact response 
of the fuselage, cabin floor, cabin furnishings (including standard and modified seats), and anthropomorphic test dummies. The 
test was conducted to simulate the vertical velocity component of a severe but survivable crash impact. A low-wing, 19-passenger 
fuselage was dropped from a height of 11 ' 2" resulting in a vertical impact velocity of 26.8 ft/sec. The airframe was configured 
to simulate a typical flight condition, including seats (normal and experimental), simulated occupants, and cargo. For the test the 

32 



wings were removed; the vertical and horizontal stabilizers were removed; the landing gear was removed; and the pilot and copilot 
seats were not installed. The data collected in the test and future tests will supplement the existing basis for improved seat and 
restraint systems for commuter category 14 Code of Federal Regulation (CFR) Part 23 airplanes. The test article was fully instru- 
mented with accelerometers and load cells. Seventy-nine data channels were recorded. Results of the test are as follows: - the fuse- 
lage experienced an impact in the range of 149-160 g's, with an impact pulse duration of 9-10 milliseconds - the simulated 
occupants experienced g levels in the range of 32-45 g's with a pulse duration of 44-61 milliseconds - the test was considered to 
be a severe but definitely survivable impact - the fuselage structure maintained a habitable environment during and after the impact 
- the seat tracks remained attached to the fuselage along the entire length of the fuselage - all standard seats remained in their tracks 
after the impact - all exits remained operable 
DTIC 

Fuselages; Crashes; Damage Assessment; Impact Tests; Drop Tests; Airframes; Transport Aircraft; Wings; Aircraft 
Compartments 

19980227362 NASA Dryden Flight Research Center, Edwards, CA USA 

Measurements Obtained During the First Landing of the North American X-15 Research Airplane 

McKay, James M., NASA Dryden Flight Research Center, USA; Oct. 1959; 38p; In English 

Report No.(s): NASA-TM-X-207; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The first landing of the X-15 airplane was made at 8:43 a.m., June 8, 1959, on the hard surface of Rogers Dry Lake. One 
purpose of the first-glide flight was to evaluate the effectiveness of the landing-gear system. Some results are presented of the 
landing-approach characteristics, the impact period, and the runout phase of the landing maneuver. The results indicate that the 
touchdown was accomplished at a vertical velocity of 2.0 feet per second for the main gear and 13.5 feet per second for the nose 
gear. These vertical velocities were within the values of sinking speeds established by structural design limitations. However, per- 
manent structural deformation occurred in the main-landing-gear system as a result of the landing, and a reevaluation of the gear 
is being made by the manufacturer. The landing occurred at a true ground speed of 158 knots for main-gear touchdown at an angle 
of attack of 8.50. The incremental acceleration at the main gear was 2.7g and 7.39 at the nose gear as a result of the landing. The 
incremental acceleration at the center of gravity of the airplane was 0.6g for the main-gear impact and 2.4g for the nose-gear 
impact. The incremental acceleration at the main gear as a result of the nose-gear impact was 4.8g. The extreme rearward location 
of the main-gear skids appears to offer satisfactory directional stability characteristics during the run- out phase of the landing. 
No evidence of nosewheel shimmy was indicated during the impact and runout phase of the landing despite the absence of a 
shimmy damper on the nose gear. The maximum amount of skid wear as a result of the landing was on the order of 0.005 inch. 
No appreciable amount of tire wear was indicated for the dual, corotating nosewheels. 
Author 

X-15 Aircraft; Structural Design; Directional Stability; Center of Gravity; Aerodynamic Characteristics; Angle of Attack; 
Deformation 

19980227405 NASA Langley Research Center, Hampton, VA USA 

Free-Flight Investigation of Radio Controlled Models with Parawings 

Hewes, Donald E., NASA Langley Research Center, USA; Sep. 1961; 34p; In English 

Report No.(s): NASA-TN-D-927; L-1374; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A free-flight investigation of two radio-controlled models with parawings, a glider configuration and an airplane (powered) 
configuration, was made to evaluate the performance, stability, and methods of controlling parawing vehicles. The flight tests 
showed that the models were stable and could be controlled either by shifting the center of gravity or by using conventional eleva- 
tor and rudder control surfaces. Static wind-tunnel force-test data were also obtained. 
Author 

Parawings; Flight Tests; Free Flight; Wind Tunnel Tests; Aerodynamic Stability; Gliders; Aerodynamic Configurations; Aircraft 
Stability 

19980227410 NASA Langley Research Center, Hampton, VA USA 

Investigation of Low-Subsonic Flight Characteristics of a Model of a Hypersonic Boost-Glide Configuration Having a 78 

deg. Delta Wing 

Paulson, John W., NASA Langley Research Center, USA; Shanks, Robert E., NASA Langley Research Center, USA; May 1961; 

32p; In English 

Report No.(s): NASA-TN-D-894; L-452; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

33 



An investigation of the low-subsonic stability and control characteristics of a model of a hypersonic boost-glide configuration 
having 78 deg. sweep of the leading edge has been made in the Langley full-scale tunnel. The model was flown over an angle-of-at- 
tack range from 10 to 35 deg. Static and dynamic force tests were made in the Langley free-flight tunnel. The investigation showed 
that the longitudinal stability and control characteristics were generally satisfactory with neutral or positive static longitudinal 
stability. The addition of artificial pitch damping resulted in satisfactory longitudinal characteristics being obtained with large 
amounts of static instability. The most rearward center-of-gravity position for which sustained flights could be made either with 
or without pitch damper corresponded to the calculated maneuver point. The lateral stability and control characteristics were satis- 
factory up to about 15 deg. angle of attack. The damping of the Dutch roll oscillation decreased with increasing angle of attack; 
the oscillation was about neutrally stable at 20 deg. angle of attack and unstable at angles of attack of about 25 deg. and above. 
Artificial damping in roll greatly improved the lateral characteristics and resulted in flights being made up to 35 deg. angle of 
attack. 
Author 
Flight Characteristics; Dynamic Stability; Boostglide Vehicles; Delta Wings; Swept Wings; Free Flight; Aircraft Control 

19981)227431 NASA Langley Research Center, Hampton, VA USA 

Flight Tests of a 1/6-Scale Model of the Hawker P 1127 Jet VTOL Airplane 

Smith, Charles C, Jr., NASA Langley Research Center, USA; 1961; 144p; In English 

Report No.(s): NASA-TM-SX-531; L-1484; No Copyright; Avail: CASI; A07, Hardcopy; A02, Microfiche 

An experimental investigation has been made to determine the dynamic stability and control characteristics of a 1/6-scale 
flying model of the Hawker P HP7 jet vertical-take-off -and-landing (VTOL) airplane in hovering and transition flight. The model 
was powered by a counter-rotating ducted fan driven by compressed-air jets at the tips of the fan blades. In hovering flight the 
model was controlled by jet-reaction controls which consisted of yaw and pitch jets at the extremities of the fuselage and a roll 
jet on each wing tip. In forward flight the model was controlled by conventional ailerons and rudder and an all-movable horizontal 
tail. In hovering flight the model could be flown smoothly and easily, but the roll control was considered too weak for rapid maneu- 
vering or hovering in gusty air. Transitions from hovering to normal forward flight and back to hovering could be made smoothly 
and consistently and with only moderate changes in longitudinal trim. The model had a static longitudinal instability or pitch-up 
tendency throughout the transition range, but the rate of divergence in the pitch-up was moderate and the model could be controlled 
easily provided the angle of attack was not allowed to become too high. In both the transition and normal forward flight conditions 
the lateral motions of the model were difficult to control at high angles of attack, apparently because of low directional stability 
at small angles of sideslip. The longitudinal stability of the model in normal forward flight was generally satisfactory, but there 
was a decided pitch-up tendency for the flap-down condition at high angles of attack. In the VTOL landing approach condition, 
with the jets directed straight down or slightly forward, the nose-down pitch trim required was greater than in the transitions from 
hovering to forward flight, but the longitudinal instability was about the same. Take-offs and landings in still air could be made 
smoothly although there was a slight unfavorable ground effect on lift and a nose-down change in pitch trim near the ground. Short 
take-offs and landings could be made smoothly and consistently although the model experienced a decided nose-up change in 
pitching moment as it climbed out of ground effect. 
Author 

Flight Tests; Dynamic Stability; Vertical Takeoff Aircraft ; Directional Stability; Scale Models; Hovering; Horizontal Flight; 
Aerodynamic Characteristics 

19980227443 Naval Postgraduate School, Monterey, CA USA 

Performance Enhancements to Joint Army /Navy Rotorcraft Analysis and Design (JANRAD) Software and Graphical. 

User Interface (GUI) 

Hucke, William L., Naval Postgraduate School, USA; Jun. 1998; 360p; In English 

Report No.(s): AD-A350646; No Copyright; Avail: CASI; A16, Hardcopy; A03, Microfiche 

The Joint Army/Navy Rotorcraft Analysis and Design (JANRAD) computer program was developed at the Naval Postgradu- 
ate School to perform performance, stability and control, and rotor dynamics analysis during preliminary helicopter design efforts. 
This thesis is the continuation of a previous work in which a Graphical User Interface (GUI) was developed and implemented as 
the front end of the NPS program. Due to the complexity of the GUI design, only the Performance module of JANRAD was com- 
pleted by the prior student. This thesis expands the capabilities of the Performance module, and the JANRAD code, by adding 
graphical output of performance results, improved rotor sizing capabilities, resources for user defined blade elements and non-lin- 
ear blade twist, airfoil meshing capabilities, and additional reference airfoil data corrected for compressibility effects. It contains 
the basic architecture for the Stability and Control module GUI. Additionally, utilizing actual UH-60A Black Hawk airfoil and 
test flight data as inputs, JANRAD version 5.0 was run to validate its output with the test flight results, and those produced in a 

34 



prior thesis by JANRAD version 3.1 (1995). Excellent agreement was demonstrated in all flight regimes. Utilizing airfoil data 
corrected for compressibility effects, high altitude runs resulted in much better correlation with test flight results than those experi- 
enced in 1995 using uncorrected airfoil data. A JANRAD Users Guide was updated and is included in Appendix A. 
DTIC 

Graphical User Interface; Computer Programs; Aircraft Design; Software Engineering; Helicopters 

19980227452 NASA Langley Research Center, Hampton, VA USA 

Incipient- and Developed-Spin and Recovery Characteristics of a Modern High-Speed Fighter Design with Low Aspect 

Ratio as Determined from Dynamic-Model Tests 

Lee, Henry A., NASA Langley Research Center, USA; Libbey, Charles E., NASA Langley Research Center, USA; Dec. 1961; 

22p; In English 

Report No.(s): NASA-TN-D-956; L-1662; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Incipient- and developed-spin and recovery characteristics of a modern high-speed fighter design with low aspect ratio have 
been investigated by means of dynamic model tests. A 1/7-scale radio-controlled model was tested by means of drop tests from 
a helicopter. Several 1/25-scale models with various configuration changes were tested in the Langley 20-foot free-spinning tun- 
nel. Model results indicated that generally it would be difficult to obtain a developed spin with a corresponding airplane and that 
either the airplane would recover of its own accord from any poststall motion or the poststall motion could be readily terminated 
by proper control technique. On occasion, however, the results indicated that if a post-stall motion were allowed to continue, a 
fully developed spin might be obtainable from which recovery could range from rapid to no recovery at all, even when optimum 
control technique was used. Satisfactory recoveries could be obtained with a proper-size tail parachute or strake, application of 
pitching-, rolling-, or yawing-moment rockets, or sufficient differential deflection of the horizontal tail. 
Author 

Fighter Aircraft; Low Aspect Ratio; Aircraft Design; Wind Tunnel Tests; Scale Models; Stability Derivatives; Dynamic Stability; 
Aerodynamic Stability; Drop Tests 

19980227737 NASA Langley Research Center, Hampton, VA USA 

Analysis of Effects of Interceptor Roll Performance and Maneuverability on Success of Collision-Course Attack 

Phillips, William H., NASA Langley Research Center, USA; Aug. 1961; 42p; In English 

Report No.(s): NASA-TN-D-952; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An attempt has been made to determine the importance of rolling performance and other factors in the design of an interceptor 
which uses collision-course tactics. A graphical method is presented for simple visualization of attack situations, by means of 
diagrams showing vectoring limits, that is, the ranges of interceptor position and heading from which attacks may be successfully 
completed, the relative importance of rolling performance and normal-acceleration capability in determining the success of 
attacks is illustrated. The results indicate that the reduction in success of attacks due to reduced rolling performance (within the 
limits generally acceptable from the pilots' standpoint) is very small, whereas the benefits due to substantially increasing the nor- 
mal-acceleration capability are large. Additional brief analyses show that the optimum speed for initiating a head-on attack is often 
that corresponding to the upper left-hand corner of the V-g diagram. In these cases, increasing speed beyond this point for given 
values of normal acceleration and radar range rapidly decreases the width of the region from which successful attacks can be initi- 
ated. On the other hand, if the radar range is increased with a variation somewhere between the first and second power of the inter- 
ceptor speed, the linear dimensions of the region from which successful attacks can be initiated vary as the square of the interceptor 
speed. 
Author 
Roll; Lateral Control; Maneuverability ; Interceptors; Fighter Aircraft; Collisions; Aircraft Performance 

19980227749 NASA Langley Research Center, Hampton, VA USA 

Some Landing Studies Pertinent to Glider-Reentry Vehicles 

Houbolt, John C, NASA Langley Research Center, USA; Batterson, Sidney A., NASA Langley Research Center, USA; Aug. 

1960; 24p; In English 

Report No.(s): NASA-TN-D-448; L-1066; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Results are presented of some landing studies that may serve as guidelines in the consideration of landing problems of glider- 
reentry configurations. The effect of the initial conditions of sinking velocity, angle of attack, and pitch rate on impact severity 
and the effect of locating the rear gear in various positions are discussed. Some information is included regarding the influence 

35 



of landing-gear location on effective masses. Preliminary experimental results on the slideout phase of landing include sliding 

and rolling friction coefficients that have been determined from tests of various skids and all-metal wheels. 

Author 

Gliders; Reentry Vehicles; Spacecraft Landing; Spacecraft Reentry 

19980227754 NASA Langley Research Center, Hampton, VA USA 

Effect of Blade Cutout on Power Required by Helicopters Operating at High Tip-Speed Ratios 

Gessow, Alfred, NASA Langley Research Center, USA; Gustafson, F. B., NASA Langley Research Center, USA; Sep. 1960; 20p; 

In English 

Report No.(s): NASA-TN-D-382; L-696; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A numerical study was made of the effects of blade cutout on the power required by a sample helicopter rotor traveling at 
tip-speed ratios of 0.3, 0.4, and 0.5. The amount of cutout varied from to 0.5 of the rotor radius and the calculations were carried 
out for a thrust coefficient-solidity ratio of 0.04. In these calculations the blade within the cutout radius was assumed to have zero 
chord. The effect of such cutout on profile-drag power ranged from almost no effect at a tip-speed ratio of 0.3 to as much as a 60 
percent reduction at a tip-speed ratio of 0.5. Optimum cutout was about 0.3 of the rotor radius. Part of the large power reduction 
at a tip-speed ratio of 0.5 resulted from a reduction in tip-region stall, brought about by cutout. For tip-speed ratios greater than 
0.3, cutout also effected a significant increase in the ability of the rotor to overcome helicopter parasite drag. It is thus seen that 
the adverse trends (at high tip-speed ratios) indicated by the uniform-chord theoretical charts are caused in large measure by the 
center portion of the rotor. The extent to which a modified-design rotor can actually be made more efficient at high speeds than 
a uniform-chord rotor will depend in practice on the degree of success in minimizing the blade plan form near the center and on 
special modifications in center-section profiles. A few suggestions and estimates in regard to such modifications are included 
herein. 
Author 
Helicopters; High Speed; Tip Speed; Rotor Blades (Turbomachinery); Thrust; Openings 

19980227756 NASA Langley Research Center, Hampton, VA USA 

A Flight Study of the Coo version Maneuver of a Tilt-Duct VTOL Aircraft 

Tapscott, Robert J., NASA Langley Research Center, USA; Kelley, Henry L., NASA Langley Research Center, USA; Nov. 1960; 

14p; In English 

Report No.(s): NASA-TN-D-372; L-891; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Flight records are presented from an early flight test of a wing-tip mounted tilting-ducted-fan, vertical-take-off and landing 
(VTOL) aircraft configuration. Time histories of the aircraft motions, control positions, and duct pitching-moment variation are 
presented to illustrate the characteristics of the aircraft in hovering, in conversion from hovering to forward flight, and in conver- 
sion from forward flight to hovering. The results indicate that during essentially continuous slow level- flight conversions, this 
aircraft experiences excessive longitudinal trim changes. Studies have shown that the large trim changes are caused primarily by 
the variation of aerodynamic moments acting on the duct units. Action of the duct-induced downwash on the horizontal stabilizer 
during the conversion also contributes to the longitudinal trim variations. Time histories of hovering and slow vertical descent 
in the final stages of landing in calm air show angular motions of the aircraft as great as +/- 10 deg. about all axes. Stick and pedal 
displacements required to control the aircraft during the landing maneuver were on the order of 50 to 60 percent of the total travel 
available. 
Author 
Vertical Takeoff Aircraft; Aircraft Configurations; Aircraft Control; Flight Tests; Ducted Fans; Flight Characteristics 

19980227770 NASA Langley Research Center, Hampton, VA USA 

Preliminary Investigation of a Paraglider 

Rogallo, Francis M., NASA Langley Research Center, USA; Lowry, John G., NASA Langley Research Center, USA; Croom, 

Delwin R., NASA Langley Research Center, USA; Taylor, Robert T., NASA Langley Research Center, USA; Aug. 1960; 28p; 

In English 

Report No.(s): NASA-TN-D-443; L-827; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A preliminary investigation of the aerodynamic and control characteristics of a flexible glider similar to a parachute in 
construction has been made at the Langley Research Center to evaluate its capabilities as a reentry glider. Preliminary weight esti- 
mates of the proposed vehicle indicate that such a structure can be made with extremely low wing loading. Maximum temperatures 
during the reentry maneuver might be held as low as about 1,500 F. The results of wind-tunnel and free-glide tests show that the 
glider when constructed of nonporous material performed extremely well at subsonic speeds and could be flown at angles of attack 

36 



from about 200 to 900. At supersonic speeds the wing showed none of the unfavorable tendencies exhibited by conventional para- 
chutes at these speeds, such as squidding and breathing. Several methods of packing and deploying the glider have been success- 
fully demonstrated. The results of this study indicate that this flexible-lifting-surface concept may provide a lightweight 
controllable paraglider for manned space vehicles. 
Author 
Lifting Reentry Vehicles; Wind Tunnel Tests; Reentry; Parachutes; Paraglider s; Supersonic Speed; Manned Spacecraft 

19980227777 NASA Langley Research Center, Hampton, VA USA 

Data from a Static-Thrust Investigation of Lstnge-Scaie General Research VTOL-STOL Model in Ground Effect 

Huston, Robert J., NASA Langley Research Center, USA; Winston, Matthew M., NASA Langley Research Center, USA; Aug. 

1960; 66p; In English 

Report No.(s): NASA-TN-D-397; L-987; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

The model was tested at two different elevations with the wing pivot at 1.008 and 2.425 propeller diameters above the ground. 
The slipstream of the propellers was deflected by tilting the wing and propellers, by deflections of large-chord trailing-edge flaps, 
and by combinations of flap deflection and wing tilt. Tests were conducted over a range of propeller disk loadings from 7.41 to 
29.70 pounds per square foot. Force data for the complete model and pressure distributions for the wing and flaps behind one 
propeller were recorded and are presented in tabular form without analysis. 
Author 

Vertical Takeoff Aircraft; Aircraft Configurations; Body-Wing Configurations; Aircraft Design; Short Takeoff Aircraft; Aircraft 
Structures; Flaps (Control Surfaces); Rotor Aerodynamics 

19980227804 NASA Langley Research Center, Hampton, VA USA 

Longitudinal Aerodynamic Characteristics of a Four-Propeller Deflected Slipstream VTOL Model Including the Effects 

of Ground Proximity 

Kuhn, Richard E., NASA Langley Research Center, USA; Grunwald, Kalman J., NASA Langley Research Center, USA; Nov. 

1960; 136p; In English 

Report No.(s): NASA-TN-D-248; L-735; No Copyright; Avail: CASI; A07, Hardcopy; A02, Microfiche 

Results are presented of a wind-tunnel investigation of the longitudinal stability, control, and performance characteristics of 
a model of a four-propeller deflected-slipstream VTOL airplane in the transition speed range. These results indicate that steady 
level-flight transition and descending flight-path angles up to 7 or 8 deg. out of the region of ground effect can be accomplished 
without wing stall being encountered. In general, the pitching moments out of ground proximity can be adequately trimmed by 
programming the stabilizer incidence to increase with increasing flap deflection, except for a relatively large diving moment in 
the hovering condition. The deflection of the slipstream onto the horizontal tail in proximity of the ground substantially increases 
the diving moment in hovering, unless the tail is set at a large nosedown incidence. 
Author 

Aerodynamic Characteristics; Longitudinal Stability; Wind Tunnel Tests; Longitudinal Control; Vertical Takeoff Aircraft; Propel- 
ler Slipstreams; Flight Paths; Flapping 

06 
AIRCRAFT INSTRUMENTATION 

Includes cockpit and cabin display devices; and flight instruments. 



Naval Postgraduate School, Monterey, CA USA 
?stem Development for a Rotary Wing Unmanned Aerial Vehicle 
Greer, Daniel S., Naval Postgraduate School, USA; Jun. 1998; 122p; In English 
Report No.(s): AD-A350437; No Copyright; Avail: CASI; A06, Hardcopy; A02, Microfiche 

The Naval Postgraduate School has developed a successful Rapid Flight Test Prototyping System (RFTPS) for the develop- 
ment of software for remote computer control of fixed wing Unmanned Aerial Vehicles (UAV). This thesis reviews the work 
accomplished to mount sensors on a small remote controlled helicopter with instrumentation compatible with the RFTPS: an iner- 
tial measurement unit, a Global Positioning System (GPS) receiver, an altitude sensor and associated power supply and telemetry 
equipment. A helicopter with sufficient lift capability was selected and a lightweight aluminum structure was built to serve as both 
an avionics platform for the necessary equipment and also as a landing skid. Since the altitude sensors used for fixed wing UAV's, 
such as barometric sensors and GPS, do not provide sufficient accuracy for low altitude hover control, a lightweight, precision 

37 



altimeter was developed using ultrasound technology. Circuitry was developed to drive a Polaroid 6500 Series Ranging Module 
and process the output data in a form compatible with the RFTPS avionics architecture. Flight testing revealed severe vibrations 
throughout the helicopter. An alternative avionics package of reduced size was constructed to house the sonic altimeter and a three- 
axis accelerometer. Subsequent test flight results and recommendations for further research are provided. 
DTIC 
Avionics; Rotary Wings; Fixed Wings; Flight Tests 

19981)227412 NASA Langley Research Center, Hampton, VA USA 

A Simulator Steely of the Effectiveness of a Pilot's Indicator which Combined Angle of Attack and Rate of Change of Total 

Pressure as Applied to the Take-Off Rotation and Climbout of a Supersonic Transport 

Hall, Albert W., NASA Langley Research Center, USA; Harris, Jack E., NASA Langley Research Center, USA; Sep. 1961; 24p; 

In English 

Report No.(s): NASA-TN-D-948; L-1644; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A simulator study has been made to determine the effectiveness of a single instrument presentation as an aid to the pilot in 
controlling both rotation and climbout path in take-off. The instrument was basically an angle-of-attack indicator, biased with a 
total-pressure-rate input as a means of suppressing the phugoid oscillation. Linearized six-degree-of-freedom equations of motion 
were utilized in simulating a hypothetical supersonic transport as the test vehicle. Each of several experienced pilots performed 
a number of simulated take-offs, using conventional flight instruments and either an angle-of-attack instrument or the combined 
angle-of-attack and total-pressure-rate instrument. The pilots were able to rotate the airplane, with satisfactory precision, to the 
15 deg. angle of attack required for lift-off when using either an angle-of-attack instrument or the instrument which combined 
total-pressure-rate with angle of attack. At least 4 to 6 second-S appeared to be required for rotation to prevent overshoot, particu- 
larly with the latter instrument. The flight paths resulting from take-offs with simulated engine failures were relatively smooth 
and repeatable within a reasonably narrow band when the combined angle-of-attack and total-pressure-rate instrument presenta- 
tion was used. Some of the flight paths resulting from take-offs with the same engine-failure conditions were very oscillatory when 
conventional instruments and an angle-of-attack instrument were used. The pilots considered the combined angle-of-attack and 
total- pressure-rate instrument a very effective aid. Even though they could, with sufficient practice, perform satisfactory clim- 
bouts after simulated engine failure by monitoring the conventional instruments and making correction based on their readings, 
it was much easier to maintain a smooth flight path with the single combined angle-of-attack and total-pressure-rate instrument. 
Author 

Indicating Instruments; Flight Instruments; Angle of Attack; Takeoff; Simulators; Aircraft Pilots; Supersonic Transports; Pres- 
sure Distribution 

19980227734 NASA Langley Research Center, Hampton, VA USA 

Repeatability, Drift, and Aftereffect of Three Types of Aircraft Altimeters 

Gracey, William, NASA Langley Research Center, USA; Stell, Richard E., NASA Langley Research Center, USA; Jul. 1961; 42p; 

In English 

Report No.(s): NASA-TN-D-922; L-1580; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

In a series of laboratory tests of a number of sensitive altimeters 5 (Air Force type C-12 and C-13) and of precision altimeters 
(Air Force 8 type MA- 1), the repeatability was determined for the full range of each type of instrument the drift characteristics 
were determined during 1 -hour periods at various altitudes, and the drift and aftereffect were measured for a variety of simulated 
flights representative of some civil and military operations. For comparable altitude ranges, the repeatability errors of the C-12 
and C-13 types were generally of the same order while those of the MA-1 type were somewhat smaller. The drift and aftereffect 
of the C-12 instruments were smaller than those of the C-13 instruments, and the drift and aftereffect of the MA-1 altimeters were 
considerably smaller than those of both types of the sensitive instruments. The drift of each of the three types of altimeters was 
found to increase with altitude and the drift of the precision type was found to increase with increasing rate of altitude change 
preceding the drift test. 
Author 
Altimeters; Drift (Instrumentation); Aircraft Instruments; Altimetry 

19980227799 NASA Langley Research Center, Hampton, VA USA 

Repeatability of the Over- Ail Errors of as Airplane Altimeter Installation in Landing-Approach Operations 

Gracey, William, NASA Langley Research Center, USA; Stickle, Joseph W., NASA Langley Research Center, USA; May 1961; 

22p; In English 

Report No.(s): NASA-TN-D-898; L-1333; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

38 



Flight tests have been conducted to determine the repeatability of the over-all altimetry errors in the landing-approach condi- 
tion of two sensitive altimeters (Air Force type C-12) installed in the cockpit of a transport airplane and of four precision altimeters 
(Air Force type MA-1) installed in a photo -observer. Data were obtained through a speed range of 62 to 100 knots during 42 land- 
ing-approach operations conducted on four different days. The results of the tests show that the repeatability errors of the two 
sensitive altimeters are +/- 35 feet and +/- 39 feet. These errors are of the same order as the maximum repeatability error measured 
in previous tests of eleven airplanes of the same type. For each of the four flights of the present tests the mean values of the data 
obtained with the two sensitive altimeters shifted by relatively large amounts, apparently because of the interaction of the stability 
and aftereffect- recovery characteristics of the instruments. For concurrent measurements of the over-all errors of the four preci- 
sion altimeters, it is concluded that for comparable installations, the repeatability errors measured with these altimeters would be 
smaller than those measured with the sensitive altimeters. 
Author 
Altimeters; Transport Aircraft; Altimetry; Flight Tests; Aircraft Instruments; Landing Instruments; Installing 

07 
AIRCRAFT PROPULSION AND POWER 

Includes prime propulsion systems and systems components, e.g., gas turbine engines and compressors; and onboard auxiliary 
power plants for aircraft. 

19980221787 Naval Air Warfare Center, Weapons Div., China Lake, CA USA 

FAA TS3-L-13L Turbine Fragment Containment Test Final Report 

Frankenberger, C. E., Ill, Naval Air Warfare Center, USA; Jun. 1998; 24p; In English 

Contract(s)/Grant(s):DTFA03-95-X-90019 

Report No.(s): PB98-159965; DOT/FAA/AR-98/22; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The result of the FAA T53-L-13L engine turbine disk fragment containment test is presented in this report. A containment 
ring was fabricated with a 0.014 inch titanium inner and outer sleeve. One-inch-thick Kevlar 29 ballistic fabric made up the pri- 
mary structure of the containment ring. The ring was reinforced with titanium rods inserted through the fabric and laser welded 
to the inner and outer sleeves. The engine and containment ring were installed in an UH-1 Hey helicopter. The second stage power 
turbine disk was notched so that the disk would rupture at approximately 20,400 rpm. The engine was started and immediately 
accelerated to minimize the chance of a premature rupture. The event was recorded on high-speed film at 4000 pictures per second. 
The disk ruptured as the engine accelerated through 19,629 rpm. The disk ruptured into three equal section (approximately 3.6 
lbs. each). The result was a contained tri-hub burst with minor bulging of the containment ring and little sign of distress to the 
airframe. This test demonstrated the capability to contain a tri-hub burst on a medium sized turboshaft helicopter engine. 
NTIS 
Turbines; Fragments ; Containment; Turbine Engines 

19980223961 NASA Dryden Flight Research Center, Edwards, CA USA 
Flight Testing the Linear Aerospike SR-71 Experiment (LASRE) 

Corda, Stephen, NASA Dryden Flight Research Center, USA; Neal, Bradford A., NASA Dryden Flight Research Center, USA; 
Moes, Timothy R., NASA Dryden Flight Research Center, USA; Cox, Timothy H., NASA Dryden Flight Research Center, USA; 
Monaghan, Richard C, NASA Dryden Flight Research Center, USA; Voelker, Leonard S., NASA Dryden Flight Research Center, 
USA; Corpening, Griffin P., NASA Dryden Flight Research Center, USA; Larson, Richard R., NASA Dryden Flight Research 
Center, USA; Powers, Bruce G., Analytical Services and Materials, Inc., USA; Sep. 1998; 24p; In English; 30th, 15-17 Sep. 1998, 
Reno, NV, USA; Sponsored by Society of Flight Test Engineers, USA 
Contract(s)/Grant(s): RTOP 242-33-02-00-23 

Report No.(s): NASA/TM- 1998-206567; H-2280; NAS 1.15:206567; No Copyright; Avail: CASI; A03, Hardcopy; A01, Micro- 
fiche 

The design of the next generation of space access vehicles has led to a unique flight test that blends the space and flight 
research worlds. The new space vehicle designs, such as the X-33 vehicle and Reusable Launch Vehicle (RLV), are powered by 
linear aerospike rocket engines. Conceived of in the 1960's, these aerospike engines have yet to be flown, and many questions 
remain regarding aerospike engine performance and efficiency in flight, to provide some of these data before flying on the X-33 
vehicle and the RLV, a spacecraft rocket engine has been flight-tested atop the NASA SR-71 aircraft as the Linear Aerospike SR-71 
Experiment (LASRE). A 20 percent-scale, semispan model of the X-33 vehicle, the aerospike engine, and all the required fuel 
and oxidizer tanks and propellant feed systems have been mounted atop the SR-71 airplane for this experiment. A major technical 

39 



objective of the LASRE flight test is to obtain installed-engine performance flight data for comparison to wind-tunnel results and 
for the development of computational fluid dynamics-based design methodologies. The ultimate goal of firing the aerospike 
rocket engine in flight is still forthcoming. An extensive design and development phase of the experiment hardware has been com- 
pleted, including approximately 40 ground tests. Five flights of the LASRE and firing the rocket engine using inert liquid nitrogen 
and helium in place of liquid oxygen and hydrogen have been successfully completed. 
Author 

SR-71 Aircraft; X-33 Reusable Launch Vehicle; Reusable Launch Vehicles; Flight Tests; Computational Fluid Dynamics; Aero- 
spike Engines; Aerodynamic Characteristics 



19980223991) NASA Lewis Research Center, Cleveland, OH USA 

Effects of Tip Clearance and Casing Recess on Heal Transfer and Stage Efficiency in Axial Turbines 

Ameri, A. A., AYT Corp., USA; Steinthorsson, E., NASA Lewis Research Center, USA; Rigby, David L., NYMA, Inc., USA; 

Aug. 1998; 15p; In English; Turbo, 2-5 Jun. 1998, Stockholm, Sweden; Sponsored by American Society of Mechanical Engineers, 

USA 

Contract(s)/Grant(s): NAS3-27571; RTOP 523-26-13-00 

Report No.(s): NASA/CR-1998-208514; E-11287; NAS 1.26:208514; ICOMP-98-04; No Copyright; Avail: CASI; A03, Hardco- 

py; A01, Microfiche 

Calculations were performed to assess the effect of the tip leakage flow on the rate of heat transfer to blade, blade tip and 
casing. The effect on exit angle and efficiency was also examined. Passage geometries with and without casing recess were consid- 
ered. The geometry and the flow conditions of the GE-E 3 first stage turbine, which represents a modem gas turbine blade were 
used for the analysis. Clearance heights of 0%, 1%, 1.5% and 3% of the passage height were considered. For the two largest clear- 
ance heights considered, different recess depths were studied. There was an increase in the thermal load on all the heat transfer 
surfaces considered due to enlargement of the clearance gap. Introduction of recessed casing resulted in a drop in the rate of heat 
transfer on the pressure side but the picture on the suction side was found to be more complex for the smaller tip clearance height 
considered. For the larger tip clearance height the effect of casing recess was an orderly reduction in the suction side heat transfer 
as the casing recess height was increased. There was a marked reduction of heat load and peak values on the blade tip upon 
introduction of casing recess, however only a small reduction was observed on the casing itself. It was reconfirmed that there is 
a linear relationship between the efficiency and the tip gap height. It was also observed that the recess casing has a small effect 
on the efficiency but can have a moderating effect on the flow underturning at smaller tip clearances. 
Author 
Gas Turbines; Heat Transfer; Axial Flow Turbines; Blade Tips; Clearances; Recesses 



19980227187 NASA Lewis Research Center, Cleveland, OH USA 

Investigation of the Effects of Low Reynolds Number Operation on the Performance < 



Forrette, Robert E., NASA Lewis Research Center, USA; Holeski, Donald E., NASA Lewis Research Center, USA; Plohr, Henry 

W., NASA Lewis Research Center, USA; Sep. 1959; 52p; In English 

Report No.(s): NASA-TM-X-9; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

High-altitude turbojet performance is adversely affected by the effects of low air density. This performance loss is evaluated 
as a Reynolds number effect, which represents the increased significance of high fluid viscous forces in relation to dynamic fluid 
forces as the Reynolds number is decreased. An analytical and experimental investigation of the effects of low Reynolds number 
operation on a single-stage, high-work-output turbine with a downstream stator was carried out at Reynolds numbers of 182,500, 
39,600, and 23,000, based on average rotor-design flow conditions. At low Reynolds numbers and turbulent flow conditions, 
increased viscous losses caused decreased effective flow area, and thus decreased weight flow, torque, and over-all efficiency at 
a given equivalent speed and pressure ratio. Decreasing the Reynolds number from 182,500 to 23,000 at design equivalent speed 
resulted in a 5.00-point loss in peak over-all turbine efficiency for both theory and experiment. The choking equivalent weight 
flow decreased 2.30 percent for these conditions. Limiting loading work output was reached at design equivalent speed for all 
three Reynolds numbers. The value of limiting loading work output at design speed decreased 4.00 percent as Reynolds number 
was decreased from 182,500 to 23,000. A theoretical performance-prediction method using basic boundary-layer relations gave 
good agreement with experimental results over most of the performance range at a given Reynolds number if the experimental 
and analytical design operating conditions were carefully matched at the highest Reynolds number with regard to design perfor- 

40 



mance parameters. High viscous losses in the inlet stator and rotor prevented the attainment of design equivalent work output at 

the lowest Reynolds number of 23,000. 

Author 

Low Reynolds Number; Performance Prediction; Turbulent Flow; Turbojet Engines; Turbines; Design Analysis; Engine Design; 
Engine Parts 

19980227324 Naval Air Warfare Center, Weapons Div., China Lake, CA USA 

FA A T53-L-13L Turbine Fragment Containment Test Final Report 

Frankenberger, C. E.; Jun. 1998; 16p; In English 

Contract(s)/Grant(s):DTFA03-95-X-90019 

Report No.(s): AD-A350454; DOT/FAA/AR-98/22; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The result of the FAA T53-L-13L engine turbine disk fragment containment test is presented in this report. A containment 
ring was designed and fabricated by Pepin Associates, Inc. and provided to the Naval Air Warfare Center, Weapons Division by 
the William J. Hughes Technical Center. This ring was fabricated with a 0.014-inch titanium inner and outer sleeve. One-inch- 
thick Kevlar 29 ballistic fabric made up the primary structure of the containment ring. The ring was reinforced with titanium rods 
inserted through the fabric and laser welded to the inner and outer sleeves. The engine and containment ring were installed in an 
UH-1 Huey helicopter. The second stage power turbine disk was notched so that the disk would rupture at approximately 20,400 
rpm. The engine was started and immediately accelerated to minimize the chance of a premature rupture. The event was recorded 
on high-speed film at 4000 pictures per second. The disk ruptured as the engine accelerated through 19,629 rpm. The disk ruptured 
into three equal sections (approximately 3.6 lbs. each). The result was a contained tri-hub burst with minor bulging of the contain- 
ment ring and little sign of distress to the airframe. This test demonstrated the capability to contain a tri-hub burst on a medium 
sized turboshaft helicopter engine. 
DTIC 
Helicopter Engines; Turbines; Rings; Fragments; Containment; Turboshafts 

08 
AIRCRAFT STABILITY AND CONTROL 

Includes aircraft handling qualities; piloting; flight controls; and autopilots. 

19981)223075 NASA Langley Research Center, Hampton, VA USA 

Exploratory Investigation at Mach Number of 2.01 of the Longitudinal Stability and Control Characteristics of a Winged 

Reentry Configu ration 

Foster, Gerald V., NASA Langley Research Center, USA; Dec. 1959; 22p; In English 

Report No.(s): NASA-TM-X-178; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An investigation has been conducted to determine the longitudinal stability and control characteristics of a reentry configura- 
tion at a Mach number of 2.01. The configuration consisted of clipped delta wing with hinged wing-tip panels. The results indicate 
that deflecting the wing-tip panels from a position normal to the wing chord plane to a position coincident with the wing chord 
plane resulted in a stabilizing change in the pitching-moment characteristics but did not significantly affect the nonlinearity of 
the pitching-moment variation with angle of attack. The trailing-edge controls were effective in producing pitching moment 
throughout the angle-of-attack range for control deflections up to at least 600. The control deflection required for trim, however, 
varied nonlinearly with angle of attack. It would appear that this nonlinearity as well as the maximum deflection required for trim 
could be greatly decreased by utilizing a leading-edge control in conjunction with a trailing-edge control. 
Author 

Reentry Vehicles; Delta Wings; Longitudinal Stability; Supersonic Speed; Aerodynamic Stability; Spacecraft Stability; Wind Tun- 
nel Stability Tests; Spacecraft Control 

19981)223588 NASA Langley Research Center, Hampton, VA USA 



Jones, George W., Jr., NASA Langley Research Center, USA; Farmer, Moses G., NASA Langley Research Center, USA; 1959; 

54p; In English 

Report No.(s): NASA-TM-SX-242; L-648; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

A transonic flutter investigation has been made of models of the T-tail of the Blackburn NA-39 airplane. The models were 
dynamically and elastically scaled from measured airplane data in accordance with criteria which include a flutter safety margin. 

41 



The investigation was made in the Langley transonic blowdown tunnel and covered a Mach number range from 0.73 to 1.09 at 
simulated altitudes extending to below sea level. The results of the investigation indicated that, if differences between the mea- 
sured model and scaled airplane properties are disregarded, the airplane with the normal value of stabilizer pitching stiffness 
should have a stiffness margin of safety of at least 32 percent at all Mach numbers and altitudes within the flight boundary. How- 
ever, the airplane with the emergency value of stabilizer pitching stiffness would not have the required margin of safety from sym- 
metrical flutter at Mach numbers greater than about 0.85 at low altitudes. First-order corrections for some differences between 
the measured model and scaled airplane properties indicated that the airplane with the normal value of stabilizer pitching stiffness 
would still have an adequate margin of safety from flutter and that the flutter safety margin for the airplane with the emergency 
value of stabilizer pitching stiffness would be changed from inadequate to adequate. However, the validity of the corrections is 
questionable. 
Author 
Transonic Flutter; Wind Tunnel Tests; Wind Tunnel Models; Tail Assemblies; Flutter Analysis; Aeroelasticity 



19980223607 NASA Langley Research Center, Hampton, VA USA 

The Lateral Response of Airplanes to Random Atmospheric Turbulence 

Eggleston, John M., NASA Langley Research Center, USA; Phillips, William H., NASA Langley Research Center, USA; 1960; 

62p; In English 

Report No.(s): NASA-TR-R-74; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

Random variations of gust velocities across the span and along the fuselage are considered. In part 1 a simplified method is 
presented in which the gust velocities are represented as rolling gusts, yawing gusts, and side gusts. A sample calculation proce- 
dure is presented for obtaining the response of the airplane in each degree of freedom. 
Author 
Atmospheric Turbulence; Gusts; Prediction Analysis Techniques; Yaw 



19981)223619 NASA Ames Research Center, Moffett Field, CA USA 

An Examination of Handling Qualities Criteria for V/STOL Aircraft 

Anderson, Seth B., NASA Ames Research Center, USA; Jul. 1960; 56p; In English 

Report No.(s): NASA-TN-D-331; A-406; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

A study has been undertaken to define hand-ling qualities criteria for V/STOL aircraft. With the current military requirements 
for helicopters and airplanes as a framework, modifications and additions were made for conversion to a preliminary set of 
V/STOL requirements using a broad background of flight experience and pilots' comments from VTOL and STOL aircraft, BLC 
(boundary-layer-control) equipped aircraft, variable stability aircraft, flight simulators and landing approach studies. The report 
contains a discussion of the reasoning behind and the sources of information leading to suggested requirements. The results of 
the study indicate that the majority of V/STOL requirements can be defined by modifications to the helicopter and/or airplane 
requirements by appropriate definition of reference speeds. Areas where a requirement is included but where the information is 
felt to be inadequate to establish a firm quantitative requirement include the following: Control power and damping relationships 
about all axes for various sizes and types of aircraft; control power, sensitivity, d-amping and response for height control; dynamic 
longitudinal and dynamic lateral- directional stability in the transition region, including emergency operation; hovering steadi- 
ness; acceleration and deceleration in transition; descent rates and flight-path angles in steep approaches, and thrust margin for 
approach. 
Author 
Aircraft Control; V/STOL Aircraft; Quality; Aircraft Stability; Control Stability; Controllability; Aerodynamics 



i25 NASA Langley Research Center, Hampton, VA USA 

and Analysis of Horizontal-Tail Contribution to Longitudinal Stability of* Swept-Wing Airplanes at Low Speeds 
Neely, Robert H., NASA Langley Research Center, USA; Griner, Roland F, NASA Langley Research Center, USA; 1959; 96p; 
In English 
Report No.(s): NASA-TR-R-49; No Copyright; Avail: CASI; A05, Hardcopy; A01, Microfiche 

Air-flow characteristics behind wings and wing-body combinations are described and are related to the downwash at specific 
tall locations for unseparated and separated flow conditions. The effects of various parameters and control devices on the air- flow 
characteristics and tail contribution are analyzed and demonstrated. An attempt has been made to summarize certain data by empir- 

42 



ical correlation or theoretical means in a form useful for design. The experimental data herein were obtained mostly at Reynolds 

numbers greater than 4 x 10(exp 6) and at Mach numbers less than 0.25. 

Author 

Body-Wing Configurations; Longitudinal Stability; Swept Wings; Separated Flow; Control Equipment 

19980223923 Naval Air Warfare Center, Aircraft Div., Patuxent River, MD USA 

Nonlinear Adaptive Flight Control with a Backstepping Design Approach 

Steinberg, Marc L., Naval Air Warfare Center, USA; Page, Anthony B., Naval Air Warfare Center, USA; Jan. 1998; 12p; In 

English 

Report No.(s): AD-A350986; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

This paper examines the use of adaptive backstepping for multi-axis control of a high performance aircraft. The control law 
is demonstrated on a 6 Degree-of -Freedom simulation with nonlinear aerodynamic and engine models, actuator models with satu- 
ration, and turbulence. Simulation results are demonstrated for large pitch-roll maneuvers, and for maneuvers with failure of the 
right stabilator. There are substantial differences between the control law design and simulation models, which are used to demon- 
strate some robustness aspects of this control law. Actuator saturation is shown to be a considerable problem for this type of con- 
troller. However, the flexibility of the backstepping design provides opportunities for improvement. In particular, the Lyapunov 
function is modified so that the growth of integrated error and the rate of change of parameter growth are both reduced when the 
surface commands are growing at a rate that will likely saturate the actuators. In addition, the deadzone technique from robust 
linear adaptive control is applied to improve robustness to turbulence. 
DTIC 
Control Theory; Controllers; Supersonic Aircraft; Aircraft Models; Flight Control; Adaptive Control 

19980223967 NASA Langley Research Center, Hampton, VA USA 

Performance, Stability, and Control Investigation at Mach Numbers from 0.60 to 1.05 of a Model of the "Swallow" with 

Outer Wing Panels Swept 75 degree with and without Power Simulations 

Schmeer, James W., NASA Langley Research Center, USA; Cassetti, Marlowe D., NASA Langley Research Center, USA; Jun. 

23, 1960; 66p; In English 

Report No.(s): NASA-TM-SX-306; L-1014; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

An investigation of the performance, stability, and control characteristics of a variable-sweep arrow-wing model with the 
outer wing panels swept 75 deg. has been conducted in the Langley 16-foot transonic tunnel. Four outboard engines located above 
and below the wing provided propulsive thrust, and, by deflecting in the pitch direction and rotating in the lateral plane, also pro- 
duced control forces. The engine nacelles incorporated swept lateral and vertical fins for aerodynamic stability and control. Jet-off 
data were obtained with flow-through nacelles, simulating inlet flow; jet thrust and hot-jet interference effects were obtained with 
faired-nose nacelles housing hydrogen peroxide gas generators. Six-component force and moment data were obtained at Mach 
numbers from 0.60 to 1.05 through a range of angles of attack and angles of side-slip. Control characteristics were obtained by 
deflecting the nacelle-fin combinations as elevators, rudders, and ailerons at several fixed angles for each control. The results indi- 
cate that the basic wing-body configuration becomes neutrally stable or unstable at a lift coefficient of 0. 1 5; addition of nacelles 
with fins delayed instability to a lift coefficient of 0.30. Addition of nacelles to the wing-body configuration increased minimum 
drag from 0.0058 to 0.0100 at a Mach number of 0.60 and from 0.0080 to 0.0190 at a Mach number of 1.05 with corresponding 
reductions in maximum lift-drag ratio of 12 percent and 33 percent, respectively. The nacelle-fin combinations were ineffective 
as longitudinal controls but were adequate as directional and lateral controls. The model with nacelles and fins was directionally 
and laterally stable; the stability generally increased with increasing lift. Jet interference effects on stability and control character- 
istics were small but the adverse effects on drag were greater than would be expected for isolated nacelles. 
Author 

Aerodynamic Stability; Body-Wing Configurations; Aerodynamic Coefficients; Nacelles; Arrow Wings; Control Surfaces; Direc- 
tional Control; Transonic Speed 

19980223987 Naval Air Warfare Center, Aircraft Div., Patuxent River, MD USA 

Robust Command Augmentation System Desigo Using Genetic Methods 

Sweriduk, G. D., Optimal Synthesis, USA; Menon, P. K., Optimal Synthesis, USA; Stienberg, M. L., Naval Air Warfare Center, 

USA; Jan. 1998; 9p; In English 

Report No.(s): AD-A350849; No Copyright; Avail: CASI; A02, Hardcopy; A01, Microfiche 

This paper describes the use of a genetic search method in the design of a command augmentation system for a high-perfor- 
mance aircraft. A genetic algorithm is used in the design of H(infinity) controllers for the longitudinal and lateral-directional chan- 

43 



nels by selecting the weighting functions. The integral of absolute value of error between the actual response and that of an ideal 

model is used as the fitness criterion, along with additional terms to penalize for cross-coupling between Ps and ny; non-minimum 

phase behavior, and the closed-loop infinity-norm bound, gamma. Starting from an initial population of weighting functions, the 

algorithm generates new functions with the goal of improving the fitness. These controllers are then evaluated in a 6 degree-of- 

freedom nonlinear model of the aircraft. 

DTIC 

Genetic Algorithms; Feedback Control; Control Systems Design; Computer Aided Design; Control Theory; Flight Control 



19989227976 NASA Lewis Research Center, Cleveland, OH USA 

Static Stability and Control of Canard Configurations al Mach Numbers from 0.70 to 2.22 - Triangular Wing and € 

with Twin Vertical Tails 

Peterson, Victor L., NASA Lewis Research Center, USA; Jun. 1961; 40p; In English 

Report No.(s): NASA-TN-D-1033; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The static aerodynamic characteristics of a canard airplane configuration having twin vertical stabilizing surfaces are pre- 
sented. The model consisted of a wing and canard both of triangular plan form and aspect ratio 2 mounted on a Sears-Haack body 
of fineness ratio 12.5 and two swept and tapered wing-mounted vertical tails of aspect ratio 1.35. Data are presented for Mach 
numbers from 0.70 to 2.22 and for angles of attack from -6 to +18 deg. at and 5 deg. sideslip. Tests were made with the canard 
off and with the canard on. Nominal canard deflection angles ranged from to 10 deg. The Reynolds number was 3.68 x 10(exp 
6) based on the wing mean aerodynamic chord. Selected portions of the data obtained in this investigation are compared with pre- 
viously published results for the same model having a single vertical tail instead of twin vertical tails. Without the canard, the 
directional stability at supersonic Mach numbers and high angles of attack was improved slightly by replacing the single tail with 
twin tails. However, at a Mach number of 0.70, the directional stability of the twin-tail model deteriorated rapidly with increasing 
angle of attack above 10 deg. and fell considerably below the level for the single-tail model. At subsonic speeds the directional 
stability of the twin-tail model with the canard was comparable to that for the single-tail model and at supersonic speed it was 
considerably greater at high angles of attack. Unlike the single-tail model, the twin-tail model at 50 sideslip exhibited an unstable 
break in the variation of pitching-moment coefficient with lift coefficient near 10 deg. angle of attack for 0.70 Mach number. 
Author 

Static Aerodynamic Characteristics; Canard Configurations; Directional Stability; Static Stability; Subsonic Speed; Tail Assem- 
blies; Aerodynamic Coefficients; Delta Wings 



19989227982 NASA Ames Research Center, Moffett Field, CA USA 

The Effect of Lateral-Directional Control Coupling on Pilot Control of an Airplane as Determined in Flight and in a Fixed- 
Base Flight Simulator 

Vomaske, Richard R, NASA Ames Research Center, USA; Sadoff, Melvin, NASA Ames Research Center, USA; Drinkwater, 
Fred J., Ill, NASA Ames Research Center, USA; Nov. 1961; 46p; In English 
Report No.(s): NASA-TN-D-1141; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A flight and fixed-base simulator study was made of the effects of aileron-induced yaw on pilot opinion of aircraft lateral-di- 
rectional controllability characteristics. A wide range of adverse and favorable aileron-induced yaw was investigated in flight at 
several levels of Dutch-roll damping. The flight results indicated that the optimum values of aileron- induced yaw differed only 
slightly from zero for Dutch-roll damping from satisfactory to marginally controllable levels. It was also shown that each range 
of values of aileron-induced yawing moment considered satisfactory, acceptable, or controllable increased with an increase in the 
Dutch- roll damping. The increase was most marked for marginally controllable configurations exhibiting favorable aileron-in- 
duced yaw. Comparison of fixed-base flight simulator results with flight results showed agreement, indicating that absence of 
kinesthetic motion cues did not markedly affect the pilots' evaluation of the type of control problem considered in this study. The 
results of the flight study were recast in terms of several parameters which were considered to have an important effect on pilot 
opinion of lateral-directional handling qualities, including the effects of control coupling. Results of brief tests with a three-axis 
side-arm controller indicated that for control coupling problems associated with highly favorable yaw and cross-control tech- 
niques, use of the three-axis controller resulted in a deterioration of control relative to results obtained with the conventional center 
stick and rudder pedals. 
Author 

Directional Control; Lateral Control; Flight Simulators; Controllability; Pilot Induced Oscillation; Aircraft Stability; Aircraft 
Control 

44 



19980227088 NASA Langley Research Center, Hampton, VA USA 

Dynamic Longitudinal and Directional Stability Derivatives for a 45 deg. Sweptfoack-Wing Airplane Model at Transonic 

Speeds 

Bielat, Ralph P., NASA Langley Research Center, USA; Wiley, Harleth G., NASA Langley Research Center, USA; Aug. 1959; 

54p; In English 

Report No.(s): NASA-TM-X-39; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

An investigation was made at transonic speeds to determine some of the dynamic stability derivatives of a 45 deg. sweptback- 
wing airplane model. The model was sting mounted and was rigidly forced to perform a single-degree-of -freedom angular oscilla- 
tion in pitch or yaw of +/- 2 deg. The investigation was made for angles of attack alpha, from -4 deg. to 14 deg. throughout most 
of the transonic speed range for values of reduced-frequency parameter from 0.015 to 0.040 based on wing mean aerodynamic 
chord and from 0.04 to 0.14 based on wing span. The results show that reduced frequency had only a small effect on the damping- 
in-pitch derivative and the oscillatory longitudinal stability derivative for all Mach numbers M and angles of attack with the excep- 
tion of the values of damping coefficient near M = 1 .03 and alpha = 8 deg. to 14 deg. In this region, the damping coefficient changed 
rapidly with reduced frequency and negative values of damping coefficient were measured at low values of reduced frequency. 
This abrupt variation of pitch damping with reduced frequency was a characteristic of the complete model or wing -body-vertical- 
tail combination. The damping-in-pitch derivative varied considerably with alpha and M for the horizontal-tail-on and horizontal- 
tail-off configurations, and the damping was relatively high at angles of attack corresponding to the onset of pitch-up for both 
configurations. The damping-in-yaw derivative was generally independent of reduced frequency and M at alpha = -4 deg. to 4 
deg. At alpha = 8 deg. to 14 deg., the damping derivative increased with an increase in reduced frequency and alpha for the configu- 
rations having the wing, whereas the damping derivative was either independent of or decreased with increase in reduced fre- 
quency for the configuration without the wing. The oscillatory directional stability derivative for all configurations generally 
decreased with an increase in the reduced-frequency parameter, and, in some instances, unstable values were measured for the 
model configuration with the horizontal tail removed. 
Author 

Aircraft Models; Directional Stability; Longitudinal Stability; Transonic Speed; Wind Tunnel Tests; Sweptback Wings; Aerody- 
namic Configurations; Aircraft, Control 

19980227090 NASA Ames Research Center, Moffett Field, CA USA 

A Study of Longitudinal Control Problems at Low and Negative Damping and Stability with Emphasis on Effects of 

Motion Cues 

Sadoff, Melvin, NASA Ames Research Center, USA; McFadden, Norman M., NASA Ames Research Center, USA; Heinle, 

Donovan R., NASA Ames Research Center, USA; Jan. 1961; 54p; In English; No Copyright; Avail: CASI; A04, Hardcopy; A01, 

Microfiche 

As part of a general investigation to determine the effects of simulator motions on pilot opinion and task performance over 
a wide range of vehicle longitudinal dynamics, a cooperative NAS A-AMAL program was conducted on the centrifuge at Johns- 
ville, Pennsylvania. The test parameters and measurements for this program duplicated those of earlier studies made at Ames 
Research Center with a variable-stability airplane and with a pitch-roll chair flight simulator. Particular emphasis was placed on 
the minimum basic damping and stability the pilots would accept and on the minimum dynamics they considered controllable in 
the event of stability-augmentation system failure. Results of the centrifuge-simulator program indicated that small positive 
damping was required by the pilots over most of the frequency range covered for configurations rated acceptable for emergency 
conditions only (e.g., failure of a pitch damper). It was shown that the pilot's tolerance for unstable dynamics was dependent pri- 
marily on the value of damping. For configurations rated acceptable for emergency operation only, the allowable instability and 
damping corresponded to a divergence time to double amplitude of about 1 second. Comparisons were made of centrifuge, pitch- 
chair and fixed-cockpit simulator tests with flight tests. Pilot ratings indicated that the effects of incomplete or spurious motion 
cues provided by these three modes of simulation were important only for high-frequency, lightly damped dynamics or unstable, 
moderately damped dynamics. The pitch- chair simulation, which provided accurate angular-acceleration cues to the pilot, 
compared most favorably with flight. For the centrifuge simulation, which furnished accurate normal accelerations but spurious 
pitching and longitudinal accelerations, there was a deterioration of pilots' opinion relative to flight results. Results of simulator 
studies with an analog pilot replacing the human pilot illustrated the adaptive capability of human pilots in coping with the wide 
range of vehicle dynamics and the control problems covered in this study. It was shown that pilot-response characteristics, deduced 
by the analog-pilot method, could be related to pilot opinion. Possible application of these results for predicting flight-control 
problems was illustrated by means of an example control-problem analysis. The results of a brief evaluation of a pencil-type side- 
arm controller in the centrifuge showed a considerable improvement in the pilots' ability to cope with high-frequency, low-damp- 
ing dynamics, compared to results obtained with the center stick. This improvement with the pencil controller was attributed 

45 



primarily to a marked reduction in the adverse effects of large and exaggerated pitching and longitudinal accelerations on pilot 

control precision. 

Author 

Longitudinal Control; Cockpit Simulators; Flight Simulators; Pilot Performance; Flight Control; Dynamic Control; Cues; Flight 
Tests; Centrifuges 

19980227095 NASA Langley Research Center, Hampton, VA USA 

Flight Investigation of an Automatic Pitchup Control 

Hurt, George J., Jr., NASA Langley Research Center, USA; Whitten, James B., NASA Langley Research Center, USA; Aug. 1960; 

30p; In English 

Report No.(s): NASA-TN-D-114; L-679; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A flight investigation of an automatic pitchup control has been conducted by the National Aeronautics and Space Administra- 
tion at the Langley Research Center. The pitching-moment characteristics of a transonic fighter airplane which was subject to 
pitchup were altered by driving the stabilizer in accordance with a signal that was a function of a combination of the measured 
angle of attack and the pitching velocity. An angle-of-attack threshold control was used to preset the angle of attack at which the 
automatic pitchup-control system would begin to drive the stabilizer. No threshold control as such existed for the pitching-velocity 
signal. A summing linkage in series with the pilot's longitudinal control allowed the automatic pitchup-control system to drive 
the stabilizer 13.5 percent of the total stabilizer travel independently of the pilot's control. Tests were made at an altitude of 35,000 
feet over a Mach number range of 0.80 to 0.90. Various gearings between the control and the sensing devices were investigated. 
The automatic system was capable of extending the region of positive stability for the test airplane to angles of attack above the 
basic-airplane pitchup threshold angle of attack. In most cases a limit-cycle oscillation about the airplane pitch axis occurred. 
Author 
Pitching Moments; Fighter Aircraft; Automatic Control; Stability Tests; Longitudinal Control; Aircraft Design; Transonic Flight 

19980227100 NASA Langley Research Center, Hampton, VA USA 

Ad Experimental Investigation of the Effects of Mach Number, Stabilizer Dihedral, and Fin Torsional Stiffness on the 

Transonic Flutter Characteristics of a Tee-Tail 

Land, Norman S., NASA Langley Research Center, USA; Fox, Annie G., NASA Langley Research Center, USA; Oct. 1961; 26p; 

In English 

Report No.(s): NASA-TN-D-924; L-1611; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A transonic flutter investigation was made of elastically and dynamically scaled models of the tee-tail of a patrol bomber. 
It was found that removal of the 15 deg. dihedral of the stabilizer used on the airplane raised the flutter boundary to higher dynamic 
pressures. The effect of Mach number on the flutter boundary was different for dihedral angles of and 15 deg. The dynamic 
pressure at the flutter boundary increased approximately linearly with the torsional stiffness of the fin. High-speed motion pictures 
indicated that the flutter mode consisted primarily of fin bending and fin torsion. 
Author 

Transonic Flutter; Flutter Analysis; Tail, Assemblies; Bomber Aircraft; Mach Number; T Tail, Surfaces; Horizontal, Tail Surfaces; 
Stabilizers (Fluid Dynamics); Fins 

19980227183 NASA Langley Research Center, Hampton, VA USA 

Transonic and Supersonic Flutter Investigation of 1/2-Size Models of All-Movable Canard Surface of an Expendable Pow- 
ered Target 

Ruhlin, Charles L., NASA Langley Research Center, USA; Tuovila, W. J., NASA Langley Research Center, USA; 1961; 40p; 
In English 
Report No.(s): NASA-TM-SX-616; L-1303; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A transonic and a supersonic flutter investigation of 1/2-size models of the all -movable canard surface of an expendable pow- 
ered target has been conducted in the Langley transonic blowdown tunnel and in the Langley 9- by 18-inch supersonic aeroelastic- 
ity tunnel, respectively. The transonic investigation covered a Mach number range from 0.7 to 1.3, and the supersonic investigation 
was made at Mach numbers 1.3,2.0, and 2.55. The effects on the flutter characteristics of the models of different levels of stiffness 
and of free play in the pitch control linkage were examined. The semispan models, which were tested at an angle of attack of 
deg, had pitch springs with the scaled design and 1/2 the scaled design pitch stiffness and total free play in pitch ranging from 
to 1 deg. An additional model configuration which had a pitch spring 1/4 the scaled design pitch stiffness and no free play in pitch 
was included in the supersonic tests. All model configurations investigated were flutter free up to dynamic pressures 32 percent 

46 



greater than those required for flight throughout the Mach number range. Several model configurations were tested to considerably 

higher dynamic pressures without obtaining flutter at both transonic and supersonic speeds. 

Author 

Flutter Analysis; Wind Tunnel Tests; Semi span Models; Wind Tunnel Models; Canard. Configurations; Transonic Speed; Super- 
sonic Speed; Aeroelasticity; Aircraft Structures 



19980227306 NASA Ames Research Center, Moffett Field, CA USA 

A Self-Adaptive Missile Guidance System for Statistical Inputs 

Peery, H. Rodney, NASA Ames Research Center, USA; Nov. 1960; 34p; In English 

Report No.(s): NASA-TN-D-343; A-400; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A method of designing a self-adaptive missile guidance system is presented. The system inputs are assumed to be known in 
a statistical sense only. Newton's modified Wiener theory is utilized in the design of the system and to establish the performance 
criterion. The missile is assumed to be a beam rider, to have a g limiter, and to operate over a flight envelope where the open-loop 
gain varies by a factor of 20. It is shown that the percent of time that missile acceleration limiting occurs can be used effectively 
to adjust the coefficients of the Wiener filter. The result is a guidance system which adapts itself to a changing environment and 
gives essentially optimum filtering and minimum miss distance. 
Author 
Missiles; Guidance (Motion); Missile Control; Design Analysis; Trajectory Control 



19980227331 West Virginia Univ., Morgantown, WV USA 

Neara! Network Autopilot System for a Mathematical Model of the Boeing 747 

Cottrill, Gerald C; Aug. 04, 1998; 158p; In English 

Report No.(s): AD-A350857; Rept-98-042; No Copyright; Avail: CASI; A08, Hardcopy; A02, Microfiche 

Artificial neural networks can be defined as approximate mathematical models of the human brain's learning activities. In 
recent years neural networks have demonstrated abilities to perform autopilot and fault tolerant control tasks when applied to non- 
linear numerical aircraft simulations. Five on-line learning neural network autopilot systems, trained with the Standard and 
Extended Back-Propagation algorithms, were applied to a six degree-of -freedom non-linear simulation of a Boeing 747-200. The 
performance of the autopilots was compared based on their abilities to perform maneuvers at linear conditions and to adapt at 
non-linear conditions to restore steady state conditions. Linear maneuvers were performed by introducing reference values of alti- 
tude and speed, pitch angle, roll angle, or heading angle. The performance using the SBPA was satisfactory, but the EBPA perfor- 
mance was clearly superior throughout the entire range maneuvers while compensating for lightly damped phugoid and Dutch 
roll modes. Non-linear adaptation investigations were performed by exciting the non-linear terms in the equations of motion. The 
non-linear conditions were achieved in two ways: by simultaneously exciting pitch and roll rates with maximum elevator and aile- 
ron inputs, and the other by simultaneously exciting roll, pitch, and yaw rates with maximum elevator, aileron, and rudder inputs. 
The EBPA based controllers were able to regain steady state conditions for both non-linear tests with better transient performance 
than their SBPA counterparts. The SBPA showed only limited ability to adapt in cases where all three angular rates were excited. 
Artificial neural networks trained on-line using the Extended Back-Propagation algorithm are concluded to be better suited for 
autopilot systems for the 1/25 scale Boeing 747 based on their superior abilities to perform linear maneuvers and regain steady 
state conditions when at non-linear conditions. 
DTIC 

Automatic Pilots; Boeing 747 Aircraft; Mathematical Models; Neural Nets; On- Line Systems; Nonlinear Systems; Artificial Intel- 
ligence; Maneuvers 



f 333 NASA Langley Research Center, Hampton, VA USA 

cat Analysis of the Longitudinal Behavior of an Automatically Controlled Supersonic Interceptor During the 
Attack Phase 

Gates, Ordway B., Jr., NASA Langley Research Center, USA; Woodling, C. H., NASA Langley Research Center, USA; 1959; 
26p; In English 
Report No.(s): NASA-TR-R-19; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Theoretical analysis of the longitudinal behavior of an automatically controlled supersonic interceptor during the attack phase 
against a nonmaneuvering target is presented. Control of the interceptor's flight path is obtained by use of a pitch rate command 

47 



system. Topics lift, and pitching moment, effects of initial tracking errors, discussion of normal acceleration limited, limitations 
of control surface rate and deflection, and effects of neglecting forward velocity changes of interceptor during attack phase. 
Author 

Flight Paths; Deflection; Pitching Moments; Control Surfaces 



19980227351 NASA Langley Research Center, Hampton, VA USA 

Effects of Control-Response Characteristics on the Capability of Helicopter for Use sis a Gun Platform 

Pegg, Robert J., NASA Langley Research Center, USA; Connor, Andrew B., NASA Langley Research Center, USA; Sep. 1960; 

20p; In English 

Report No.(s): NASA-TN-D-464; L-796; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An investigation with a variable-stability helicopter was undertaken to ascertain the steadiness and ability to "hold on" to the 
target of a helicopter employed as a gun platform. Simulated tasks were per formed under differing flight conditions with the con- 
trol-response characteristics of the helicopter varied for each task. The simulated gun-platform mission included: Variations of 
headings with respect to wind, constant altitude and "swing around" to a wind heading of deg, and increases in altitude while 
performing a swing around to a wind heading of deg. The results showed that increases in control power and damping increased 
pilot ability to hold on to the target with fewer yawing oscillations and in a shorter time. The results also indicated that wind direc- 
tion must be considered in accuracy assessment. Greatest accuracy throughout these tests was achieved by aiming upwind. 
Author 
Helicopters; Wind Direction; Damping; Flight Conditions; Targets 



19980227363 NASA Langley Research Center, Hampton, VA USA 

Transonic Flutter Characteristics of a 45 cleg Sweptback Wing with Various Distributions of Ballast Along the Leading 

Edge 

Unangst, John R., NASA Langley Research Center, USA; Dec. 1959; 32p; In English 

Report No.(s): NASA-TM-X-135; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An investigation of the use of ballast at the leading edge of a sweptback wing as a flutter fix has been made. The investigation 
was conducted in the Langley transonic blowdown tunnel with wing models which had an aspect ratio of 4, sweepback of the 
quarter-chord line of 450, and a taper ratio of 0.2. Four ballast configurations, which included different amounts of ballast distrib- 
uted at two different span-wise locations, were investigated. Full-span sting-mounted models were employed. Data were obtained 
over a Mach number range from 0.65 to 1.32. Comparison of the data for the ballasted wings with data for a similar wing without 
ballast shows that in the often critical Mach number range between 0.85 and 1.05, the dynamic pressure required for flutter is 
increased by as much as 100 percent due to the addition of about 6 percent of the wing mass as ballast at the leading edge of the 
outboard sections. Furthermore, there are indications that similar benefits of leading-edge ballast can be obtained at Mach numbers 
above M = 1.1. Changing the spanwise location of the ballast and increasing the amount of the ballast by a factor of about 2 had 
very little additional effect on the dynamic pressure required for flutter. The possibility, therefore, exists that the beneficial effects 
obtained may be accomplished by using less than the minimum of about 6 percent of the wing mass as ballast as investigated in 
this paper. 
Author 
Transonic Flutter; Flutter Analysis; Critical Velocity; Sweptback Wings; Dynamic Pressure; Aspect Ratio 



19980227404 NASA Langley Research Center, Hampton, VA USA 

Low-Subsonic Measurements of the Static Stability and Control and Oscillatory Stability Derivatives of a Proposed Reen- 
try Vehicle Having an Extensible Heat Shield for High-Drag Reentry 

Johnson, Joseph L., Jr., NASA Langley Research Center, USA; Boisseau, Peter C, NASA Langley Research Center, USA; Aug. 
1961; 38p; In English 
Report No.(s): NASA-TN-D-892; L-1329; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A low-speed investigation has been made to determine the static and oscillatory longitudinal and lateral stability derivatives 
of a proposed reentry vehicle having an extensible heat shield for reentry at high angles of attack. The heat shield is extended 
forward to give the desired aerodynamic -center position for high-angle-of-attack reentry and, after completion of the reentry 
phase, is retracted to give stability and trim for gliding flight at low angles of attack. Near an angle of attack of 900 the reentry 
configuration was statically stable both longitudinally and directionally, had positive dihedral effect, and had positive damping 

48 



in roll but zero damping in yaw. The landing configuration had positive damping in pitch, roll, and yaw over the test angle-of-attack 
range but was directionally unstable and had negative dihedral effect between an angle of attack of about 10 and 20 deg. 
Author 

Longitudinal Stability; Reentry Vehicles; Static Stability; Reentry; Aerodynamic Balance; Heat Shielding; Lateral Stability; 
Subsonic Speed 



19980227742 NASA Langley Research Center, Hampton, VA USA 

Investigation of Longitudinal and Lateral Stability Characteristics of a Six-Propeller Deflected-Slips 

with Boundary-Layer Control Including Effects of Ground Proximity 

Grunwald, Kalman J., NASA Langley Research Center, USA; Jan. 1961; 94p; In English 

Report No.(s): NASA-TN-D-445; L-951; No Copyright; Avail: CASI; A05, Hardcopy; A01, Microfiche 

An investigation of the longitudinal and lateral stability and control and Performance characteristics of a six-propeller 
deflected- slipstream vertical-take-off-and-landing (VTOL) model in the transition speed range was conducted in the 17-foot test 
section of the Langley 300-MPH 7- by 10-foot tunnel. A complete analysis of the data was not conducted. A modest amount of 
blowing boundary-layer control was necessary to achieve transition without wing stall. 
Author 

Lateral Stability; Longitudinal Stability; Wind Tunnel Tests; Vertical Takeoff Aircraft; Boundary Layer Control; Propeller 
Slipstreams; Aircraft Configurations 



19981)227758 NASA Langley Research Center, Hampton, VA USA 

Stability and Control Characteristics of a Model of ao Aerial Vehicle Supported by Foui 

Parlert, Lysle P., NASA Langley Research Center, USA; Aug. 1961; 20p; In English 

Report No.(s): NASA-TN-D-937; L-1482; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The stability and control characteristics of a simple, lightly loaded model approximately one-third the size of a full-scale 
vehicle have been investigated by a series of free-flight tests. The model is representative of a type of vertically rising aircraft 
which would utilize four ducted fans as its sole source of lift and propulsion. The ducts were arranged in a rectangular pattern and 
were fixed to the airframe so that their axes of revolution were vertical for hovering flight. Control moments were provided by 
remotely controlled compressed-air jets at the sides and ends of the model. In hovering, the model in its original configuration 
exhibited divergent oscillations about both the roll and pitch axes. Because these oscillations were of a rather short period., the 
model was very difficult to control by the use of remote controls only. The model could be completely stabilized by the addition 
of a sufficient amount of artificial damping. The pitching oscillation was made easier to control by increasing the distance between 
the forward and rearward pairs of ducts. In forward flight, with the model in its original configuration, the top speed was limited 
by the development of an uncontrollable pitch-up. Large forward tilt angles were required for trim at the highest speeds attained. 
With the model rotated so that the shorter axis became the longitudinal axis, the pitch trim problem was found to be less than with 
the longer axis as the longitudinal axis. The installation of a system of vanes in the slipstream of the forward ducts reduced the 
tilt angle but increased the power required. 
Author 

Wind Tunnel Tests; Ducted Fans; Flight Characteristics; Aerodynamic Configurations; Wind Tunnel Models; Aerodynamic Sta- 
bility; Aircraft Control; Lift Fans; Vertical Takeoff 



19980227762 NASA Langley Research Center, Hampton, VA USA 

Dynamic Stability and Control Problems of Piloted Reentry from Lunar Missions 

Moul, Martin T, NASA Langley Research Center, USA; Schy, Albert A., NASA Langley Research Center, USA; Williams, James 

L., NASA Langley Research Center, USA; Nov. 1961; 22p; In English 

Report No.(s): NASA-TN-D-986; L-1764; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A fixed-base simulator investigation has been made of stability and control problems during piloted reentry from lunar mis- 
sions. Reentries were made within constraints of acceleration and skipping, in which the pilot was given simulated navigation 
tasks of altitude and heading angle commands. Vehicles considered included a blunt-face, high-drag capsule, and a low-drag lift- 
ing cone, each of which had a trim lift-drag ratio of 0.5. With the provision of three-axis automatic damping, both vehicles were 
easily controlled through reentry after a brief pilot-training period. With all dampers out, safe reentries could be made and both 

49 



vehicles were rated satisfactory for emergency operation. In damper-failure conditions resulting in inadequate Dutch roll damp- 
ing, the lifting-cone vehicle exhibited control problems due to excessive dihedral effect and oscillatory acceleration effects. 
Author 

Manned Reentry; Aerodynamic Stability; Altitude Simulation; Lateral Stability; Dynamic Stability; Lifting Reentry Vehicles; 
Lunar Exploration 

19980227771 NASA Langley Research Center, Hampton, VA USA 

Lateral Stability aocl Control Characteristics of a Four-Propeller Deflected-Slipstream VTOL Model Including the 

Effects of Ground Proximity 

Kuhn, Richard E., NASA Langley Research Center, USA; Grunwald, Kalman J., NASA Langley Research Center, USA; Jan. 

1961; 72p; In English 

Report No.(s): NASA-TN-D-444; L-895; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

The investigation of the lateral-directional stability and control characteristics of a four-propeller deflected-slipstream VTOL 
model in the transition speed range was conducted in the 17-foot test section of the Langley 300-MPH 7- by 10-foot tunnel. A 
large fairing on top of the rear fuselage was needed to eliminate directional instability in the power-off flaps-retracted condition. 
Even with this fairing some instability at small sideslip angles remained for power-on conditions with low flap deflections. The 
configuration exhibited a high level of dihedral effect which, coupled with the directional instability, will probably produce an 
undesirable Dutch roll oscillation. 
Author 

Vertical Takeoff Aircraft; Lateral Stability; Lateral Control; Wind Tunnel Tests; Wind Tunnel Models; Propeller Slipstreams; Air- 
craft Configurations 

19980227774 NASA Langley Research Center, Hampton, VA USA 

Effect al High Subsonic Speeds of Fuselage Forebody Strakes on the Static Stability and Vertical-Tail-Load Characteris- 
tics of a Complete Model Having a Delta Wing 

Polamus, Edward C, NASA Langley Research Center, USA; Spreemann, Kenneth P., NASA Langley Research Center, USA; 
May 1961; 32p; In English 
Report No.(s): NASA-TN-D-903; L-1531; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A wind-tunnel investigation at high subsonic speeds has been conducted to determine the effect of fuselage forebody strakes 
on the static stability and the vertical-tail-load characteristics of an airplane-type configuration having a delta wing. The tests were 
made at Mach numbers from 0.60 to 0.92 corresponding to Reynolds numbers from 3.0 x 10(exp 6) to 4.2 x 10(exp 6), based on 
the wing mean aerodynamic chord, and at angles of attack from approximately -2 to 24 deg. The strakes provided improvements 
in the directional stability characteristics of the wing-fuselage configuration which were reflected in the characteristics of the com- 
plete configuration in the angle-of-attack range where extreme losses in directional stability quite often occur. It was also found 
that the strakes, through their beneficial effect on the wing-fuselage directional stability, reduced the vertical-tail load per unit 
restoring moment at high angles of attack. The results also indicated that, despite the inherent tendency for strakes to produce a 
pitch-up, acceptable pitching-moment characteristics can be obtained provided the strakes are properly chosen and used in con- 
junction with a wing-body-tail configuration characterized by increasing stability with increasing lift. 
Author 

Body-Wing and Tail Configurations; Subsonic Speed; Strakes; Wind. Tunnel Tests; Wind Tunnel Models; Forebodies; Fuselages; 
Directional Stability; Aircraft Stability; Delta Wings; Static Stability 

19980227775 NASA Langley Research Center, Hampton, VA USA 
Study of an Active Control System for a Spinning Body 

Adams, J. J., NASA Langley Research Center, USA; Jun. 1961; 28p; In English 

Report No.(s): NASA-TN-D-905; L-1519; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The mission requirements for some satellites require that they spin continuously and at the same time maintain a precise direc- 
tion of the spin axis. An analog-computer study has been made of an attitude control system which is suitable for such a satellite. 
The control system provides the necessary attitude control through the use of a spinning wheel, which will provide precession 
torques, commanded by an automatic closed-loop servomechanism system. The sensors used in the control loop are rate gyro- 
scopes for damping of any wobble motion and a sun seeker for attitude control. The results of the study show that the controller 
can eliminate the wobble motion of the satellite resulting from a rectangular pulse moment disturbance and then return the spin 
axis to the reference space axis. The motion is damped to half amplitude in less than one cycle of the wobble motion. The controller 
can also reduce the motion resulting from a step change in product of inertia both by causing the new principal axis to be steadily 

50 



alined with the spin vector and by reducing the cone angle generated by the reference body axis. These methods will reduce the 

motion whether the satellite is a disk, sphere, or rod configuration. 

Author 

Active Control; Spin Reduction; Spin Dynamics; Satellite Attitude Control; Controllers; Satellite Rotation 

19980227800 NASA Langley Research Center, Hampton, VA USA 

Low-Speed Investigation of the Effects of Frequency and Amplitude of Oscillation In Sideslip on the Lateral Stability 

Derivatives of a 60 deg Delia Wing, a 45 deg Sweptback Wing and an Unswept Wing 

Lichtenstein, Jacob H., NASA Langley Research Center, USA; Williams, James L., NASA Langley Research Center, USA; May 

1961; 20p; In English 

Report No.(s): NASA-TN-D-896; L-1608; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A low-speed investigation has been conducted in the Langley stability tunnel to study the effects of frequency and amplitude 
of sideslipping motion on the lateral stability derivatives of a 60 deg. delta wing, a 45 deg. sweptback wing, and an unswept wing. 
The investigation was made for values of the reduced-frequency parameter of 0.066 and 0.218 and for a range of amplitudes from 
+/- 2 to +/- 6 deg. The results of the investigation indicated that increasing the frequency of the oscillation generally produced 
an appreciable change in magnitude of the lateral oscillatory stability derivatives in the higher angle-of- attack range. This effect 
was greatest for the 60 deg. delta wing and smallest for the unswept wing and generally resulted in a more linear variation of these 
derivatives with angle of attack. For the relatively high frequency at which the amplitude was varied, there appeared to be little 
effect on the measured derivatives as a result of the change in amplitude of the oscillation. 
Author 

Wind Tunnel Tests; Delta Wings; Sweptback Wings; Unswept Wings; Stability Derivatives; Lateral Stability; Sideslip; Wing 
Oscillations 

09 
RESEARCH AND SUPPORT FACILITIES (AIR) 

Includes airports, hangars and runways; aircraft repair and overhaul facilities; wind tunnels; shock tubes; and aircraft engine test 
stands. 

19981)223937 Federal Aviation Administration, Technical Center, Atlantic City, NJ USA 

Automated Surface Observing System (ASOS) Controller Equipment (ACE) Operational Test and Evaluation 



Horan, Colleen, Federal Aviation Administration, USA; Melillo, Michael R., Federal Aviation Administration, USA; Peio, Karen 
J., Federal Aviation Administration, USA; Nuzman, Edward F, Federal Aviation Administration, USA; Vicente, James P., Federal 
Aviation Administration, USA; May 1998; 69p; In English 
Report No.(s): AD-A350596; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

The Automated Surface Observing System (ASOS) Controller Equipment (ACE) system is a display system that provides 
weather products from the ASOS and other weather product systems to the Federal Aviation Administration (FAA) Air Traffic 
Control Towers (ATCTs), Terminal Radar Approach Control (TRACON), and other selected locations. Operational Test and Eval- 
uation (OT&E) of the ACE was conducted in four phases, commencing at the FAA William J. Hughes Technical Center in July 
1995 and concluding at the Will Rogers World Airport, Oklahoma City Oklahoma (OKC) and Dallas/Ft. Worth International Air- 
port, Irving, Texas (DFW), in April 1997. The purpose of the OT&E was to evaluate the performance of the ACE display system. 
This final report describes the results of OT&E testing conducted on the ACE. 
DTIC 
Controllers; Evaluation; Display Devices; Information Transfer 

19980223963 Veridian, Moffett Field, CA USA 

The Meal-Time Wall Interference Correction System of the NASA Ames 12-Foot Pressure Wind Tunnel 

Ulbrich, Norbert, Veridian, USA; Jul. 1998; 206p; In English 

Contract(s)/Grant(s): NAS2-13605 

Report No.(s): NASA/CR- 1998-208537; A-98-11989; NAS 1.26:208537; No Copyright; Avail: CASI; A10, Hardcopy; A03, 

Microfiche 

An improved version of the Wall Signature Method was developed to compute wall interference effects in three-dimensional 
subsonic wind tunnel testing of aircraft models in real-time. The method may be applied to a full-span or a semispan model. A 

51 



simplified singularity representation of the aircraft model is used. Fuselage, support system, propulsion simulator, and separation 
wake volume blockage effects are represented by point sources and sinks. Lifting effects are represented by semi-infinite line dou- 
blets. The singularity representation of the test article is combined with the measurement of wind tunnel test reference conditions, 
wall pressure, lift force, thrust force, pitching moment, rolling moment, and pre-computed solutions of the subsonic potential 
equation to determine first order wall interference corrections. Second order wall interference corrections for pitching and rolling 
moment coefficient are also determined. A new procedure is presented that estimates a rolling moment coefficient correction for 
wings with non-symmetric lift distribution. Experimental data obtained during the calibration of the Ames Bipod model support 
system and during tests of two semispan models mounted on an image plane in the NASA Ames 12 ft. Pressure Wind Tunnel are 
used to demonstrate the application of the wall interference correction method. 
Author 
Subsonic Flow; Support Systems; Wall Flow; Wall Pressure; Subsonic Wind Tunnels; Force Distribution 

19981)227544 Pittsburgh State Univ., KS USA 

Analysis, Specification and Implementation of an Automated Programmable Control System and Virtual Instrumenta- 
tion to Improve and Advance the Operation of the Slack Thermal High Vacuum Chamber 

Buchanan, Randy K., Pittsburgh State Univ., USA; 1997 Research Reports: NASA/ASEE Summer Faculty Fellowship Program; 
Dec. 1997, pp. 11-20; In English; Also announced as 19980227542; No Copyright; Avail: CASI; A02, Hardcopy; A03, Microfiche 
The Slack Thermal Vacuum Chamber was designed to process spacecraft and ground support equipment for the Materials 
Science Laboratory at Kennedy Space Center (KSC). The chamber recently became inoperative and was thus identified to be 
equipped with a modern control system to enable support of the launch of the Space Shuttle, expendable rockets, and their respec- 
tive payloads. Installation of a modern computerized programmable control system was performed, which included connection 
of new control hardware and complex programming of the controller. Furthermore, a virtual instrumentation system was created 
with the use of an additional computer and the incorporation of virtual instrumentation software. This report characterizes the 
evolution and successful completion of this modernization process. 
Author 
Control Systems Design; Numerical Control; Vacuum, Chambers; Programmable Logic Devices 

19980227789 NASA Langley Research Center, Hampton, VA USA 

Description of a 2-Foot Hypersonic Facility at the Langley Research Center 

Stokes, George M., NASA Langley Research Center, USA; Sep. 1961; 26p; In English 

Report No.(s): NASA-TN-D-939; L-1390; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

This report describes the mechanical and aerodynamic features of a two-foot hypersonic facility at the Langley Research Cen- 
ter. The facility provides for the testing of aerodynamic models in the Mach number range between 3 and 7 at approximate 
Reynolds numbers between 0.5 x 10(exp 6) and 1.0 x 10(exp 6). The facility was designed to obtain the needed pressure ratio 
through the use of ejector nozzles. Compressors driving the ejectors operate continuously at a pressure ratio of 4 and thus give 
the facility a continuous running capability. Curves are presented to show the ranges of total temperature, total pressure, Reynolds 
number dynamic pressure, and static pressure available in the tunnel. The flow in the test section is suitable for model tests at all 
Mach numbers between 3 and 7, although the nozzle blocks were contoured for a Mach number of 6. 
Author 
Supersonic Wind Tunnels; Aerodynamic Characteristics; Supersonic Speed; Hypersonics; Compressors 

19980227798 NASA Langley Research Center, Hampton, VA USA 

The Development of an 8-inch by 8-inch Slotted Tunnel for Mach Numbers up to 1.28 

Little, B. H., Jr., NASA Langley Research Center, USA; Cubbage, James J., Jr., NASA Langley Research Center, USA; Aug. 

1961; 14p; In English 

Report No.(s): NASA-TN-D-908; L-1005; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An 8-inch by 8-inch transonic tunnel model with test section slotted on two opposite walls was constructed in which particular 
emphasis -was given to the development of slot geometry, slot-flow reentry section, and short-diffuser configurations for good 
test-region flow and minimum total-pressure losses. Center-line static pressures through the test section, wall static pressures 
through the other parts of the tunnel, and total-pressure distributions at the inlet and exit stations of the diffuser were measured- 
With a slot length equal to two tunnel heights and 1/14 open-area-ratio slotted walls) a test region one tunnel height in length was 
obtained in which the deviation from the mean Mach number was less than +/- 0.01 up to Mach number 1.15. With 1/7 open-area- 
ratio slotted walls, a test region 0.84 tunnel heights in length with deviation less than +/- O.01 was obtained up to Mach number 
1.26. Increasing the tunnel diffuser angle from 6.4 to 10 deg. increased pressure loss through the tunnel at Mach number 1.20 from 

52 



15 percent to 20 percent of the total pressure. The use of other diffusers with equivalent angles of 10 deg. but contoured so that 
the initial diffusion angle was less than 10 deg. and the final angle was 200 reduced the losses to as low as 16 percent. A method 
for changing the test-section Mach number rapidly by controlling the flow through a bypass line from the tunnel settling chamber 
to the slot-flow plenum chamber of the test section was very effective. The test-section Mach number was reduced approximately 
5 percent in 1/8 second by bleeding into the test section a flow of air equal to 2 percent of the mainstream flow and 30 percent 
in 1/4 second with bleed flow equal to 10 percent of the mainstream flow. The rate of reduction was largely determined by the 
opening rate of the bleed-flow-control valve. 
Author 
Transonic Wind Tunnels; Slotted Wind Tunnels; Wind Tunnel Apparatus; Pressure Distribution 

10 
ASTRONAUTICS 

Includes astronautics (general); astrodynamics; ground support systems and facilities (space); launch vehicles and space vehicles; 
space transportation; space communications, spacecraft communications, command and tracking; spacecraft design, testing and 
performance; spacecraft instrumentation; and spacecraft propulsion and power. 

19980221813 Smithsonian Astrophysical Observatory, Cambridge, MA USA 

Iii~Space Transportation with Tethers Annua! Report No, 2, 1 Sep, 1997 - 31 Aug, 1998 

Lorenzini, Enrico, Smithsonian Astrophysical Observatory, USA; Estes, Robert D., Smithsonian Astrophysical Observatory, 

USA; Cosmo, Mario L., Smithsonian Astrophysical Observatory, USA; Aug. 1998; 120p; In English 

Contract(s)/Grant(s): NAG8-1303; No Copyright; Avail: CASI; A06, Hardcopy; A02, Microfiche 

The annual report covers the research conducted on the following topics related to the use of spaceborne tethers for in-space 
transportation: ProSEDS tether modeling (current collection analyses, influence of a varying tether temperature); proSEDS mis- 
sion analysis and system dynamics (tether thermal model, thermo-electro-dynamics integrated simulations); proSEDS -tether 
development and testing (tether requirements, deployment test plan, tether properties testing, deployment tests); and tethers for 
reboosting the space-based laser (mission analysis, tether system preliminary design, evaluation of attitude constraints). 
Author 
Tethering; Air Transportation; Models; Dynamic Characteristics; Manufacturing; Design Analysis 

19980223079 NASA Langley Research Center, Hampton, VA USA 

Trajectory Control for Vehicles Entering The Earth's Atmosphere at Small Flight-Path Angles 

Eggleston, John M., NASA Langley Research Center, USA; Young, John W., NASA Langley Research Center, USA; 1961; 32p; 

In English 

Report No.(s): NASA-TR-R-89; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Methods of controlling the trajectories of high-drag low -lift vehicles entering the earth's atmosphere at angles of attack near 
90 degrees and at initial entry angles up to 3 degrees are studied. The trajectories are calculated for vehicles whose angle of attack 
can be held constant at some specified value or can be perfectly controlled as a function of some measured quantity along the 
trajectory. The results might be applied in the design of automatic control systems or in the design of instruments which will give 
the human pilot sufficient information to control his trajectory properly during an atmospheric entry. Trajectory data are compared 
on the basis of the deceleration, range, angle of attack, and, in some cases, the rate of descent. The aerodynamic heat-transfer rate 
and skin temperature of a vehicle with a simple heat-sink type of structure are calculated for trajectories made with several types 
of control functions. 
Author (revised) 

Control Systems Design; Trajectory Control; Reentry; Aerodynamic Heat Transfer; Automatic Control; Aerodynamic Heating; 
Descent Trajectories 

19980223473 NASA Langley Research Center, Hampton, VA USA 

Mars Ascent Vehicle Flight Analysis 

Desai, P. N., NASA Langley Research Center, USA; Braun, R. D., NASA Langley Research Center, USA; Engelund, W. C, 

NASA Langley Research Center, USA; Cheatwood, F. M., NASA Langley Research Center, USA; Kangas, J. A., Jet Propulsion 

Lab., California Inst, of Tech., USA; 1998; lOp; In English; 7th; Thermophysics and Heat Transfer, 15-18 Jun. 1998, Albuquerque, 

NM, USA; Sponsored by American Inst, of Aeronautics and Astronautics, USA; Original contains color illustrations 

Report No.(s): AIAA Paper 98-2850; No Copyright; Avail: Issuing Activity, Hardcopy, Microfiche 

53 



The scientific objective of the Mars Surveyor Program 2005 mission is to return Mars rock, soil, and atmospheric samples 
to Earth for detailed analysis. The present investigation focuses on design of Mars Ascent Vehicle for this mission. Aerodynamic, 
aerothermodynamic, and trajectory design considerations are addressed to assess the ascent configuration, determine aerody- 
namic stability, characterize thermal protection system requirements, and ascertain the required system mass. Aerodynamic analy- 
sis reveals a subsonic static instability with the baseline configuration; however, stability augmentation options are proposed to 
mitigate this problem. The ascent aerothermodynamic environment is shown to be benign (on the order of the sea-level boiling 
point of water on Earth). As a result of these low thermal and pressure loads, a lightweight, low rigidity material can be employed 
as the aftbody aerodynamic shroud. The required nominal MAV lift-off mass is 426 kg for a December 2006 equatorial launch 
into a 300-km circular orbit with 30-degree inclination. Off-nominal aerodynamic and atmospheric conditions are shown to 
increase this liftoff mass by approximately 10%. Through performance of these analyses, the Mars Ascent Vehicle is deemed feasi- 
ble with respect to the current mission mass and size constraints. 
Author 

Planetary Geology; Mission Planning; Mars Surface; Meteorology; Design Analysis; Circular Orbits; Afterbodies; Aerothermo- 
dynamics; Aerodynamic Stability; Aerodynamic Characteristics 

19981)223479 NASA Marshall Space Flight Center, Huntsville, AL USA 
International Space Station Electrodynamic Tether Reboost Study 

Johnson, L., NASA Marshall Space Flight Center, USA; Herrmann, M., NASA Marshall Space Flight Center, USA; Jul. 1998; 
40p; In English 

Report No.(s): NASA/TM- 1998-208538; M-886; NAS 1.15:208538; No Copyright; Avail: CASI; A03, Hardcopy; A01, Micro- 
fiche 

The International Space Station (ISS) will require periodic reboost due to atmospheric aerodynamic drag. This is nominally 
achieved through the use of thruster firings by the attached Progress M spacecraft. Many Progress flights to the ISS are required 
annually. Electrodynamic tethers provide an attractive alternative in that they can provide periodic reboost or continuous drag 
cancellation using no consumables, propellant, nor conventional propulsion elements. The system could also serve as an emer- 
gency backup reboost system used only in the event resupply and reboost are delayed for some reason. 
Author 
International Space Station; Aerodynamic Drag; Tetherlines; Tethering; Drag Reduction; Orbit Decay 

19980223581 NASA Lewis Research Center, Cleveland, OH USA 

Experimental Performance of Area Ratio 200, 25 and 8 Nozzles on JP~4 Fuel and Liquid Oxygen Rocket Engine 

Lovell, J. Calvin, NASA Lewis Research Center, USA; Samanich, Nick E., NASA Lewis Research Center, USA; Barnett, Donald 

O., NASA Lewis Research Center, USA; Aug. 1960; 18p; In English 

Report No.(s): NASA-TM-X-382; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The performance of an area ratio 200 bell-shaped nozzle, an area ratio 25 bell-shaped nozzle, and an area ratio 8 conic nozzle 
on a JP-4 fuel and liquid-oxygen rocket engine has been determined. Tests were conducted using a nominal 4000-pound-thrust 
rocket in the Lewis 10- by 10-foot supersonic tunnel, which provided the altitude environment needed for fully expanded nozzle 
flow. The area ratio 200 nozzle had a vacuum thrust coefficient of 1.96, compared with 1.82 and 1.70 for the area ratio 25 and 
8 nozzles, respectively. These values are approximately equal to those for theoretical frozen expansion. The measured value of 
vacuum specific impulse for the area ratio 200 nozzle was 317 seconds for a combustion-chamber characteristic velocity of 5200 
feet per second. The vacuum-specific-impulse increase for the area-ratio increase from 8 to 200 was 46 seconds. 
Author 
Nozzle Design; Nozzle Flow; Nozzle Geometry; JP-4 Jet Fuel; Conies; Specific Impulse; Thrust 

19980223622 NASA Ames Research Center, Moffett Field, CA USA 

An Approximate Analytical Method for Studying Entry into Planetary Atmospheres 

Chapman, Dean R., NASA Ames Research Center, USA; 1959; 48p; In English 

Report No.(s): NASA-TR-R-11; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The pair of motion equations for entry into a planetary atmosphere is reduced to a single, ordinary, nonlinear differential equa- 
tion of second order by disregarding two relatively small terms and by introducing a certain mathematical transformation. The 
reduced equation includes various terms, certain of which represent the gravity force, the centrifugal acceleration, and the lift 
force. If these particular terms are disregarded, the differential equation is linear and yields precisely the solution of Allen and 
Eggers applicable to ballistic entry at relatively steep angles of descent. If all the other terms in the basic equation are disregarded 
(corresponding to negligible vertical acceleration and negligible vertical component of drag force), the resulting truncated differ- 

54 



ential equation yields the solution of Sanger for equilibrium flight of glide vehicles with relatively large lift-drag ratios. A number 
of solutions for lifting and nonlifting vehicles entering at various initial angles also have been obtained from the complete nonlin- 
ear equation. These solutions are universal in the sense that a single solution determines the motion and heating of a vehicle of 
arbitrary weight, dimensions, and shape entering an arbitrary planetary atmosphere. One solution is required for each lift-drag 
ratio. These solutions are used to study the deceleration, heating rate, and total heat absorbed for entry into Venus, Earth, Mars, 
and Jupiter. From the equations developed for heating rates, and from available information on human tolerance limits to accelera- 
tion stress, approximate conditions for minimizing the aerodynamic heating of a trimmed vehicle with constant lift-drag ratio are 
established for several types of manned entry. 
Author 
Planetary Atmospheres; Atmospheric Entry; Aerodynamic Heating; Equations of Motion; Nonlinear Equations; Design Analysis 

19980223952 NASA Ames Research Center, Moffett Field, CA USA 

A Flight Study of a Power-Off Landing Technique Applicable to Re-Entry Vehicles 

Bray, Richard S., NASA Ames Research Center, USA; Drinkwater, Fred J., NASA Ames Research Center, USA; White, Maurice 

D., NASA Ames Research Center, USA; Jul. 1960; 30p; In English 

Report No.(s): NASA-TN-D-323; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A power-off landing technique, applicable to aircraft of configurations presently being considered for manned re-entry 
vehicles, has been developed and flight tested at Ames Research Center. The flight tests used two configurations of an airplane 
for which the values of maximum lift-drag ratios were 4.0 and 2.8. Twenty -four idle-power approaches were made to an 8000-foot 
runway with touchdown point and airspeed accuracies of +/-600 feet and +/-10 knots, respectively. The landing pattern used was 
designed to provide an explicitly defined flight path for the pilot and, yet, to require no external guidance other than the pilot's 
view from the cockpit. The initial phase of the approach pattern is a constant high-speed descent from altitude aimed at a ground 
reference point short of the runway threshold. At a specified altitude and speed, a constant g pull-out is made to a shallow flight 
path along which the air-plane decelerates to the touchdown point. Repeatability and safety are inherent because of the reduced 
number of variables requiring pilot judgment, and because of the fact that a missed approach is evident at speeds and altitudes 
suitable for safe ejection. The accuracy and repeatability of the pattern are indicated by the measured results. The proposed pattern 
appears to be particularly suitable for configurations having unusual drag variations with speed in the lower speed regime, since 
the pilot is not required to control speed in the latter portions of the pattern. 
Author 
Aircraft Configurations; Reentry Vehicles; Flight Tests; Spacecraft Landing 

19980223970 NASA Langley Research Center, Hampton, VA USA 

Some Effects of Ablation Surface Roughness on the Aerodynamic Characteristics of a Reentry Vehicle at M ach Numbers 

from 030 to 1.00 

Decker, John P., NASA Langley Research Center, USA; Abel, Irving, NASA Langley Research Center, USA; Feb. 1971; 80p; 

In English 

Contract(s)/Grant(s): RTOP 124-07-17-07 

Report No.(s): NASA-TM-SX-2050; L-7288; AF-AM-833; No Copyright; Avail: CASI; A05, Hardcopy; A01, Microfiche 

A wind-tunnel investigation has been made to determine some of the effects of ablation surface roughness on the aerodynamic 
characteristics of a reentry vehicle. The vehicle selected for the investigation was the S V-5D/FV-3 PRIME vehicle which is an 
approximate 0.28-scale model of the X-24A manned low-speed flight research vehicle. The PRIME vehicle was flown on a subor- 
bital flight and subsequently retrieved by a U.S. Air Force cargo airplane. The PRIME vehicle was restored and modified for wind- 
tunnel testing and some of the effects of ablation surface roughness were determined by testing the ablated model (PRIME vehicle 
restored and modified for wind-tunnel tests) and a smooth replica. The tests were conducted at Mach numbers from 0.30 to 1.00. 
Author 

Reentry Vehicles; Suborbital Flight; Surface Roughness Effects; Ablation; Aerodynamic Characteristics; Scale Models; Wind 
Tunnel Tests; X-24 Aircraft; Aerodynamic Heating 

19980223971 NASA Langley Research Center, Hampton, VA USA 

Low Subsonic Aerodynamic Characteristics of a Reentry Spacecraft Shape with a High Hypersonic Lift-Drag Ratio 

Martin, James A., NASA Langley Research Center, USA; Decker, John P., NASA Langley Research Center, USA; Apr. 1971; 

38p; In English 

Contract(s)/Grant(s): RTOP 124-07-17-07 

Report No.(s): NASA-TM-SX-2097; L-7334; AF-AM-920; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

55 



An investigation has been conducted in the Langley low -turbulence pressure tunnel to determine the aerodynamic character- 
istics of a proposed flight-research vehicle with a delta planform and a high hypersonic lift-drag ratio. The tests were conducted 
at a Mach number of about 0.25, angles of attack from about -9 deg. to 31 deg., angles of sideslip from about -7 deg. to 12 deg., 
and Reynolds numbers based on body length from 4.5 x 10(exp 6) to 43.2 X 10(exp 6). 
Author 
Subsonic Speed; Aerodynamic Characteristics; Reentry Vehicles; Wind Tunnel Tests; Aerodynamic Coefficients 

19980223977 NASA Langley Research Center, Hampton, VA USA 
A Concept; of si Manned Satellite Reentry Which is Completed with a Glide Landing 
Cheatham, Donald C, Compiler, NASA Langley Research Center, USA; Dec. 1959; 46p; In English 
Report No.(s): NASA-TM-X-226; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A concept for a manned satellite reentry from a near space orbit and a glide landing on a normal size airfield is presented. 
The reentry vehicle configuration suitable for this concept would employ a variable geometry feature in order that the reentry 
could be made at 90 deg. angle of attack and the landing could be made with a conventional glide approach. Calculated results 
for reentry at a flight-path angle of -1 deg. show that with an accuracy of 1 percent in the impulse of a retrorocket, the desired 
flight-path angle at reentry can be controlled within 0.02 deg. and the distance traveled to the reentry point, within 100 miles. The 
reentry point is arbitrarily defined as the point at which the satellite passes through an altitude of about 70 miles. Misalignment 
of the retrorocket by 10 deg. increased these errors by as much as 0.02 deg. and 500 miles. Intra-atmospheric trajectory calculations 
show that pure drag reentries starting with flight -path angles of -1 deg. or less produce a peak deceleration of 8g. Lift created by 
varying the angle of attack between 90 and 60 deg. is effective in decreasing the maximum deceleration and allows the range to 
the "recovery" point (where transition is made from reentry to gliding flight) to be increased by as much as 2,300 miles. A sideslip 
angle of 30 deg. allows lateral displacement of the flight path by as much as 60 deg. miles. Reaction controls would provide con- 
trol-attitude alignment during the orbit phase. For the reentry phase this configuration should have low static longitudinal and roll 
stability in the 90 deg. angle-of-attack attitude. Control could be effected by leading-edge and trailing-edge flaps. Transition into 
the landing phase would be accomplished at an altitude of about 100,000 feet by unfolding the outer wing panels and pitching 
over to low angles of attack. Calculations indicate that glides can be made from the recovery point to airfields at ranges of from 
150 to 200 miles, depending upon the orientation with respect to the original course. 
Author 

Manned Reentry; Angle of Attack; Reentry Vehicles; Glide Landings; Descent Trajectories; Atmospheric Entry; Spacecraft 
Landing 

19981)223992 NASA Ames Research Center, Moffett Field, CA USA 

An Approximate Analytical Method for Studying Atmosphere Entry of Vehicles with Modulated Aerodynamic Forces 

Levy, Lionel L., Jr., NASA Ames Research Center, USA; Oct. 1960; 34p; In English 

Report No.(s): NASA-TN-D-319; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The dimensionless, transformed, nonlinear differential equation developed in NASA TR R-ll for describing the approximate 
motion and heating during entry into planetary atmospheres for constant aerodynamic coefficients and vehicle shape has been 
modified to include entries during which the aerodynamic coefficients and the vehicle shape are varied. The generality of the 
application of the original equation to vehicles of arbitrary weight, size, and shape and to arbitrary atmospheres is retained. A 
closed-form solution for the motion, heating, and the variation of drag loading parameter m/C(D)A has been obtained for the case 
of constant maximum resultant deceleration during nonlifting entries. This solution requires certain simplifying assumptions 
which do not compromise the accuracy of the results. The closed-form solution has been used to determine the variation of 
m/C(D)A required to reduce peak decelerations and to broaden the corridor for nonlifting entry into the earth's atmosphere at 
escape velocity. The attendant heating penalty is also studied. 
Author 

Atmospheric Entry; Aerospace Vehicles; Aerodynamic Coefficients; Aerodynamic Heating; Aerodynamic Configurations; Differ- 
ential Equations; Aerodynamic Forces 

19980227176 NASA Langley Research Center, Hampton, VA USA 

Structures for Reentry Heating 

Anderson, Roger A., NASA Langley Research Center, USA; Swann, Robert T., NASA Langley Research Center, USA; Sep. 

1960; 22p; In English 

Report No.(s): NASA-TM-X-313; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

56 



The basic structural approaches for dealing with reentry heating of manned vehicles are summarized. The weight and devel- 
opment status of both radiative and ablative shields are given and the application of these shields to various vehicles is indicated. 
Author 

Aerodynamic Heating; Reentry Effects; Ablation; Spacecraft Construction Materials; Manned Reentry ; Temperature Effects 



1998022721.5 NASA Langley Research Center, Hampton, VA USA 

Charts Depicting Kinematic and Heating Parameters for a Ballistic Reentry at Speeds of 26,000 to 45,000 Feet Pe 

Lovelace, Uriel M., NASA Langley Research Center, USA; Oct. 1961; 54p; In English 

Report No.(s): NASA-TN-D-968; L-1750; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

Reentry trajectories, including computations of convective and radiative stagnation-point heat transfer, have been calculated 
by using equations for a point -mass reentry vehicle entering the atmosphere of a rotating, oblate earth. Velocity was varied from 
26,000 to 45,000 feet per second; reentry angle, from the skip limit to -20 deg; ballistic drag parameter, from 50 to 200. Initial 
altitude was 400,000 feet. Explicit results are presented in charts which were computed for an initial latitude of 38 deg N and an 
azimuth of 90 deg from north. A method is presented whereby these results may be made valid for a range of initial latitude and 
azimuth angles. 
Author 
Radiative Heat Transfer; Aerodynamic Heat Transfer; Reentry Trajectories; Stagnation Point; Drag 



19980227432 NASA Langley Research Center, Hampton, VA USA 

Preliminary Results on Heat Transfer to the Afterbody of the Apollo Reentry Configuration at a M ach Number of 8 

Jones, Robert A., NASA Langley Research Center, USA; Sep. 1962; 14p; In English 

Report No.(s): NASA-TM-X-699; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Heat-transfer rates on the afterbody of the Apollo reentry configuration have been measured in a low -enthalpy wind tunnel 
at a Mach number of 8. The data have been presented as the ratio of the measured heat-transfer coefficient on the afterbody to 
the calculated heat-transfer coefficient at the stagnation point at zero angle of attack. This ratio was found to vary from a low of 
approximately 0.01 to a maximum of about 0.52 as the angle of attack varied from to 55 deg. 
Author 

Apollo Spacecraft; Hypersonic Speed; Heat Transfer Coefficients; Afterbodies; Heat Transfer; Aerothermodynamics ; Hyper- 
sonic Reentry; Wind Tunnel Tests; Reentry Effects 



19980227768 NASA Langley Research Center, Hampton, VA USA 

Flight Performance of a Spin-Stabilized 20-Inch-Diameter Solid-Propellant Spherical Rocket Motor 

Levine, Jack, NASA Langley Research Center, USA; Martz, C. William, NASA Langley Research Center, USA; Swain, Robert 

L., NASA Langley Research Center, USA; Swanson, Andrew G., NASA Langley Research Center, USA; Sep. 1960; 42p; In 

English 

Report No.(s): NASA-TN-D-441; L-596; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A successful flight test of a spin-stabilized 20-inch-diameter solid-propellant rocket motor having a propellant mass fraction 
of 0.92 has been made. The motor was fired at altitude after being boosted by a three-stage test vehicle. Analysis of the data indi- 
cates that a total impulse of 44,243 pound-second with a propellant specific impulse of approximately 185 was achieved over a 
total action time of about 12 seconds. These results are shown to be in excellent agreement with data from ground static firing 
tests of these motors. The spherical rocket motor with an 11-pound payload attained a velocity of 15,620 feet per second (m = 
16.7) with an incremental velocity increase for the spherical motor stage of 12,120 feet per second. 
Author 
Solid Propellant Rocket Engines; Flight Characteristics; Flight Tests; Propulsion System, Performance 

57 



11 

CHEMISTRY AND MATERIALS 

Includes chemistry and materials (general); composite materials; inorganic and physical chemistry; metallic materials; nonmetallic 
materials; propellants and fuels; and materials processing. 

19980227110 NASA Goddard Space Flight Center, Greenbelt, MD USA 

Algorithm for Estimating the Plume Centerlinc Temperature and Ceiling Jet Temperature in the Presence of a Hot Upper 

Layer 

Davis, William D., National Inst, of Standards and Technology, USA; Notarianni, Kathy A., National Inst, of Standards and 

Technology, USA; Tapper, Phillip Z., NASA Goddard Space Flight Center, USA; Jun. 1998; 34p; In English 

Report No.(s): PB98-146152; NISTIR-6178; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The experiments were designed to provide insight into the behavior of jet fuel fires in aircraft hangars and to study the impact 
of these fires on the design and operation of a variety of fire protection systems. As a result, the test series included small fires 
designed to investigate the operation of UV/TR detectors and smoke detectors as well as large fires which were used to investigate 
the operation of ceiling mounted heat detectors and sprinklers. The impact of the presence or absence of draft curtains was also 
studied in the 15 m hangar. It is shown that in order to predict the plume centerline temperature within experimental uncertainty, 
the entrainment of the upper layer gas must be modeled. For large fires, the impact of a changing radiation fraction must also be 
included in the calculation. The dependence of the radial temperature profile of the ceiling jet as a function of layer development 
is demonstrated and a ceiling jet temperature algorithm which includes the impact of a growing layer is developed. 
DTIC 
Plumes; Algorithms; Temperature Profiles; Hangars; Full Scale Tests; Estimating 

19980227194 NASA Lewis Research Center, Cleveland, OH USA 

Recombination of Hydrogen- Air Combustion Products in an Exhaust Nozzle 

Lezberg, Erwin A., NASA Lewis Research Center, USA; Lancashire, Richard B., NASA Lewis Research Center, USA; Aug. 

1961; 38p; In English 

Report No.(s): NASA-TN-D-1052; E-1246; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Thrust losses due to the inability of dissociated combustion gases to recombine in exhaust nozzles are of primary interest for 
evaluating the performance of hypersonic ramjets. Some results for the expansion of hydrogen-air combustion products are 
described. Combustion air was preheated up to 33000 R to simulate high-Mach-number flight conditions. Static-temperature mea- 
surements using the line reversal method and wall static pressures were used to indicate the state of the gas during expansion. 
Results indicated substantial departure from the shifting equilibrium curve beginning slightly downstream of the nozzle throat 
at stagnation pressures of 1.7 and 3.6 atmospheres. The results are compared with an approximate method for determining a freez- 
ing point using an overall rate equation for the oxidation of hydrogen. 
Author 
Combustion Products; Hydrogen; Stagnation Pressure; Static Pressure; Hypersonics; Ramjet Engines 

12 
ENGINEERING 

Includes engineering (general); communications and radar; electronics and electrical engineering; fluid mechanics and heat transfer; 
instrumentation and photography; lasers and masers; mechanical engineering; quality assurance and reliability; and structural 
mechanics. 

19981)223039 Nagasaki Univ., The Faculty of Engineering, Japan 

Characteristics of Fluid Dynamics and Noise in Laminar Flow Fans (Effects of Diameter of Impeller on Characteristics) 
Kodama, Yoshio, Nagasaki Univ., Japan; Hayashi, Hidechito, Nagasaki Univ., Japan; Tanaka, Kiyohiro, Nagasaki Univ., Japan; 
Fukui, Tomomi, Nagasaki Univ., Japan; Murahata, Kazuhiro, Nagasaki Univ., Japan; Reports of the Faculty of Engineering, Naga- 
saki University; Jan. 1993; ISSN 0286-0902; Volume 23, No. 40, pp. 17-23; In Japanese; No Copyright; Avail: Issuing Activity, 
Hardcopy, Microfiche 

The effects of six design parameters, the diameter of impeller, the rotational frequency, the gap of two disks, the number of 
disks, the clearance between casing wall and front shroud, the disk thickness on pressure coefficient were theoretically clarified 
over a wide range of fan flow rates and the scroll of casing, the diameter of impeller on the noise radiated from fan. The agreement 
between the predicted and experimental results of the pressure coefficient is satisfactory if the modified equation of velocity ratio 

58 



V(uth)/u = f and empirical equation of K(m) were used. The experimental results show that the fluid dynamic characteristics were 

improved and the sound pressure level risen by increasing the diameter of impeller. 

Author 

Turbofans; Design Analysis; Aerodynamic Noise; Laminar Flow; Flow Characteristics; Impellers; Sound Pressure 

19980223949 Nagasaki Univ., The Faculty of Engineering, Japan 

Characteristics of Fluid Dynamics and Noise in Counter-Rotating Fan 

Kodama, Yoshio, Nagasaki Univ., Japan; Hayashi, Hidechito, Nagasaki Univ., Japan; Tanaka, Kiyohiro, Nagasaki Univ., Japan; 

Yamaguti, Akihiro, Nagasaki Univ., Japan; Reports of the Faculty of Engineering, Nagasaki University; Jan. 1993; ISSN 

0286-0902; Volume 23, No. 40, pp. 9-15; In Japanese; No Copyright; Avail: Issuing Activity, Hardcopy, Microfiche 

The effects of asymmetry of the electric motor support, the distance between two impellers on the fluid dynamic characteris- 
tics and the fan noise were investigated experimentally with a counter rotating fan. Moreover, the comparison of the fan noise 
and the fluid dynamic characteristics between the counter rotating fan and a two stage rotor fan was made. It is concluded from 
these experimental results that the fan noise generated from symmetric support fan is lower 3 to 6 dB than that of asymmetric 
support fan and the fluid dynamic characteristics of the former is superior to that of the latter. The distance between two impellers 
is larger, the fan efficiency and the fan noise become lower. The fluid dynamic characteristics of the counter rotating fan with 
9-blades is superior to that of the two stage rotor fan, but the noise generated from the former is higher than that of the latter. 
Author 
Fan Blades; Electric Motors; Counter Rotation; Aerodynamic Noise; Fluid Flow 

19981)223589 NASA Langley Research Center, Hampton, VA USA 

Tire-to-Surface Friction-Coefficient Measurements with a C-123B Airplane on Various Runway Surfaces 

Sawyer, Richard H., NASA Langley Research Center, USA; Kolnick, Joseph J., NASA Langley Research Center, USA; 1959; 

36p; In English 

Report No.(s): NASA-TR-R-20; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An investigation was conducted to obtain information on the tire-to-surface friction coefficients available in aircraft braking 
during the landing run. The tests were made with a C-123B airplane on both wet and dry concrete and bituminous pavements and 
on snow-covered and ice surfaces at speeds from 12 to 115 knots. Measurements were made of the maximum (incipient skidding) 
friction coefficient, the full-skidding (locked wheel) friction coefficient, and the wheel slip ratio during braking. 
Author 
Runways; Coefficient of Friction; Pavements; Tires; C-123 Aircraft 

19980223601 NASA Langley Research Center, Hampton, VA USA 

Heat-Transfer and Pressure Measurements on a Flat-Face Cylinder at a Mach Number Range of 2.49 to 4.44 

Burbank, Paige B., NASA Langley Research Center, USA; Stallings, Robert L., Jr., NASA Langley Research Center, USA; Aug. 

1959; 24p; In English 

Report No.(s): NASA-TM-X-19; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Heat-transfer coefficients and pressure distributions were obtained on a 4-inch-diameter flat-face cylinder in the Langley 
Unitary Plan wind tunnel. The measured stagnation heat-transfer coefficient agrees well with 55 percent of the theoretical value 
predicted by the modified Sibulkin method for a hemisphere. Pressure measurements indicated the dimensionless velocity gradi- 
ent parameter r du\ a(sub t) dx, where x=0 at the stagnation point was approximately 0.3 and invariant throughout the Mach number 
range from 2.49 to 4.44 and the Reynolds number range from 0.77 x 10(exp 6) to 1.46 x 10(exp 6). The heat-transfer coefficients 
on the cylindrical afterbody could be predicted with reasonable accuracy by flat-plate theory at an angle of attack of deg. At 
angles of attack the cylindrical afterbody stagnation-line heat transfer could be computed from swept-cylinder theory for large 
distances back of the nose when the Reynolds number is based on the distance from the flow reattachment points. 
Author 

Heat Transfer; Pressure Measurement; Mach Number; Aerodynamic Heat Transfer; Flat Plates; Stagnation Point; Aerothermo- 
dynamics 

19980223606 NASA Ames Research Center, Moffett Field, CA USA 

A Study of the Simulation of Flow with Free-Stream Mach Number 1 in a Choked Wind Tunnel 

Spreiter, John R., NASA Ames Research Center, USA; Smith, Donald W., NASA Ames Research Center, USA; Hyett, B. Jeanne, 

NASA Ames Research Center, USA; 1960; 40p; In English 

Report No.(s): NASA-TR-R-73; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

59 



The degree to which experimental results obtained under choking conditions in a wind tunnel with solid walls simulate those 
associated with an unbounded flow with free-stream Mach number 1 is investigated for the cases of two-dimensional and axisym- 
metric flows. It is found that a close resemblance does indeed exist in the vicinity of the body, and that the results obtained in this 
way are generally at least as accurate as those obtained in a transonic wind tunnel with partly open test section. Some of the results 
indicate, however, that substantial interference effects may be encountered wider certain conditions, both in choked wind tunnels 
and in transonic wind tunnels, and that reduction of these interference effects to acceptable limits may require the use of models 
of unusually small size. 
Author 
Axi symmetric Flow; Transonic Speed; Test Chambers; Transonic Wind Tunnels; Free Flow 



19981)223610 NASA Lewis Research Center, Cleveland, OH USA 

Flow in the Base Region of Axisymmetric and Two-Dimensional Configi 

Beheim, Milton A., NASA Lewis Research Center, USA; 1961; 34p; In English 

Report No.(s): NASA-TR-R-77; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A theoretical and experimental investigation has been conducted of the pressure distribution on the surface of either a circular 
cylinder or a truncated cone located within the base region of another circular cylinder at Mach number 2. A similar analysis of 
pressure distribution was made for rearward-facing two-dimensional steps, and theoretical results were compared with experi- 
mental results of earlier investigations. Effects of base bleed were also studied with the axisymmetric configurations. 
Author (revised) 

Backward Facing Steps; Circular Cylinders; Supersonic Speed; Pressure Distribution; Boundary Layer Flow; Circular Cones; 
Aerodynamic Configurations 



19980223969 NASA Lewis Research Center, Cleveland, OH USA 

Effect of External Boundary Layer on Performance of Axisymmetric Inlet at Mach Numbers of 3.0 and 2,5 

Samanich, N. E., NASA Lewis Research Center, USA; Barnett, D. O., NASA Lewis Research Center, USA; Salmi, R. J., NASA 

Lewis Research Center, USA; Sep. 1959; 18p; In English 

Report No.(s): NASA-TM-X-49; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The effect of an external boundary layer on the performance of an axisymmetric external-internal-compression inlet was eval- 
uated at Mach numbers of 3.0 and 2.5 and Reynolds numbers from 2.2 to 0.5 x 10(exp 6) per foot. The inlet was tested at locations 
up to two-thirds of the way into the 1 .7- and 9.0-inch boundary layers generated by a flat plate and the tunnel floor, respectively. 
The inlet could be readily started at all conditions tested, including those where the boundary layer was separated upstream of 
the inlet by the various shock systems during the restart cycle. Although the inlet performance decreased with increasing immer- 
sion into the boundary layer at both Mach numbers, the inlet was more sensitive to boundary-layer ingestion at the design Mach 
number of 3.0. 
Author 

Aircraft Structures; Structural Design; Supersonic Speed; Boundary Layers; Internal Compression Inlets; Supersonic Inlets; Axi- 
symmetric Bodies; Wind Tunnel Tests 



19980223975 NASA Langley Research Center, Hampton, VA USA 

Elevated-Temperature Tests Under Static and Aerodynamic Conditions on Corrugated-Stiffened Pa 

Groen, Joseph M., NASA Langley Research Center, USA; Rosecrans, Richard, NASA Langley Research Center, USA; Sep. 1959; 

38p; In English 

Report No.(s): NASA-TM-X-34; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Thermal-insulating panels made of 0.005-inch-thick corrugated-stiffened sheets of Inconel X, backed by either bulk or reflec- 
tive insulation, were tested under static and aerodynamic conditions at elevated temperatures up to 1,8000 F in front of a quartz- 
tube radiant heater and in a blowdown wind tunnel at a Mach number of 1.4. The tests were performed to provide information 
on the structural integrity and insulating effectiveness of thermal-insulating panels under the effects of aerodynamic heating. 
Static radiant-heating tests showed that the bulk insulation protected a load-carrying structure better than did the reflective insula- 
tion; however, the bulk insulation was much heavier than the reflective insulation and made the panel assemblies about three times 
as thick. Three of the four panels tested in the heated supersonic wind tunnel fluttered and failed dynamically. However, one panel 

60 



demonstrated that flutter can be alleviated considerably with proper edge support. The panels deflected toward the heater (or into 

the airstream) at a rate which was primarily dependent on the temperature difference through the panel thickness. 

Author 

Wind Tunnel Tests; Aerodynamic Heating; High Temperature Tests; Insulated Structures; Static Tests; Dynamic Tests; Inconel 
(Trademark); Thermal Degradation ; Supersonic Speed 

19980223976 NASA Ames Research Center, Moffett Field, CA USA 

An Experimental Investigation of Boundary-Layer Control for Drag Reduction of a Swept- Wing Section at Low Speed 

and High Reynolds Numbers 

Gault, Donald E., NASA Ames Research Center, USA; Oct. 1960; 20p; In English 

Report No.(s): NASA-TN-D-320; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An investigation of laminar boundary-layer control by suction for purposes of drag reduction at low speed and high Reynolds 
numbers has been conducted in the Ames 12-Foot Pressure Wind Tunnel. The model was a 72.96-inch-chord wing panel, swept 
back 30 deg., which was installed between end plates to approximate a wing of infinite span. The airfoil section employed was 
a modified NACA 66-012 in the streamwise direction. Tests were limited to controlling the flow over only the upper surface of 
the model. Seventeen individually controllable suction chambers were provided below the surface to induce flow through 93 span- 
wise slots in the surface between the 0.0052- and 0.97-chord stations. Tests were made at angles of attack of deg., +/- 1 .0 deg., 
+/- 1.5 deg., and -2.0 deg. for Reynolds numbers from approximately 1.5 x 10(exp 6) to 4.0 x 10(exp 6) per foot. In general, essen- 
tially full-chord laminar flow was obtained for all conditions with small suction quantities. Minimum profile-drag coefficients 
of about 0.0005 to 0.0006 were obtained for the slotted surface at maximum values of the Reynolds number; these values include 
the Power required to induce suction as an equivalent drag. 
Author 

Boundary Layer Control; Aerodynamic Drag; Laminar Boundary Layer; Swept Wings; Wind Tunnel Tests; Suction; Wing Panels; 
Drag Reduction; End Plates 

19980227978 NASA Ames Research Center, Moffett Field, CA USA 

Forces and Moments on Sphere-Cone Bodies in Newtonian Flow 

Dickey, Robert R., NASA Ames Research Center, USA; Dec. 1961; 20p; In English 

Report No.(s): NASA-TN-D-1203; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The static longitudinal aerodynamic characteristics of a family of sphere-cone combinations (fineness ratios from 1.0 to 6.0) 
were computed by means of Newtonian impact theory. The effects of angle of attack, fineness ratio, and center-of-gravity location 
are shown. The results indicate that, with the center of gravity at or near the center of volume, the sphere-cone combinations are 
statically stable at trim points that provide low to moderate lift-drag ratios. In general, the lift-drag ratio increased with increasing 
fineness ratio. As an example, with the center of gravity at the center of volume, the lift-drag ratio at trim was increased from 
approximately 0.05 to 0.56 by increasing the fineness ratio from 1.2 to 6.0. 
Author 

Static Aerodynamic Characteristics; Circular Cones; Fineness Ratio; Newton Theory; Longitudinal, Stability; Newtonian Fluids; 
Aerodynamic Forces; Moments; Spheres 

19989227998 NASA Ames Research Center, Moffett Field, CA USA 

Full-Scale Wind-Tunnel Tests of Blowing Boundary-Layer Control Applied to a Helicopter Rotor 

McCloud, John L., Ill, NASA Ames Research Center, USA; Hall, Leo P., NASA Ames Research Center, USA; Brady, James A., 

NASA Ames Research Center, USA; Sep. 1960; 36p; In English 

Report No.(s): NASA-TN-D-335; A-380; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A full-scale wind-tunnel test was conducted of two boundary-layer-control applications to a 44-foot diameter helicopter rotor. 
Blowing from a nozzle near the leading edge of the blades delayed retreating blade stall. Results also indicated that delay of retreat- 
ing blade stall could be obtained by cyclic blowing with a lower flow rate than that required for continuous blowing. It was found 
that blowing applied through a nozzle at mid-chord had no effect on retreating blade stall. 
Author 
Boundary Layer Control; Full Scale Tests; Wind Tunnel Tests; Blowing; Leading Edges; Rotary Wings 

19980227113 NASA Lewis Research Center, Cleveland, OH USA 

Reduction of Jet Penetration in a Cross-Flow by Using Tabs 

Zaman, K. B. M. Q., NASA Lewis Research Center, USA; 1998; 8p; In English; 34th; Propulsion, 13-15 Jul. 1998, Cleveland, 

61 



OH, USA; Sponsored by American Inst, of Aeronautics and Astronautics, USA 

Report No.(s): AIAA Paper 98-3276; No Copyright; Avail: Issuing Activity, Hardcopy, Microfiche 

A tab placed suitably on a nozzle that produces a jet in a cross-flow can reduce the penetration of the jet. This effect, achieved 
when the tab is placed on the windward side of the nozzle relative to the cross flow, may be of interest in film cooling applications. 
Wind tunnel experiments are carried out, in the momentum ratio (J) range of 10-90, to investigate the tab geometry that would 
maximize this effect. The preliminary results show that a 'delta tab' having a base width approximately fifty percent of the nozzle 
diameter may be considered optimum. With a given tab size, the effect is more pronounced at higher J. Reduction in jet penetration 
by as much as 40% is observed. Comparable reduction in jet penetration is also obtained when a triangular shaped tab is placed 
flush with the tunnel wall or with its apex tilted down into the jet nozzle (the 'delta tab' being the configuration in which the apex 
is tilted up). However, the delta tab involves the least flow blockage and pressure loss. Relative to the baseline case, the lateral 
spreading of the jet is found to be more with the delta tab but less with other orientations of the tab. 
Author 
Tabs (Control Surfaces); Cross Flow; Flow Characteristics; Flow Geometry; Aerodynamic Characteristics; Nozzle Flow 

19980227181 NASA Langley Research Center, Hampton, VA USA 

An Experimental Investigation and Correlation of the Ileal Reduction to Nonporous Surfaces Behind a Porous Leading 

Edge Through Which Coolant is Ejected 

Witte, William G., NASA Langley Research Center, USA; Rashis, Bernard, NASA Langley Research Center, USA; Mar. 1960; 

34p; In English 

Report No.(s): NASA-TM-X-235; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A configuration of a wing segment having constant chord thickness, deg. sweep, a porous steel semicircular leading edge, 
and solid Inconel surfaces was tested in a Mach number 2.0 ethlyene -heated high-temperature air jet. Measurements were made 
of the wing surface temperatures at chordwise stations for several rates of helium flow through the porous leading edge. The inves- 
tigation was conducted at stagnation temperatures ranging from 500 F to 2,400 F, at Reynolds numbers per foot ranging from 0.3 
x 10(exp 7) to 1 .2 x 10(exp 7), and at angles of attack of 0, +/- 5, and +/- 15 deg. The results indicated that the reduction of wing 
surface temperatures with respect to their values for no coolant flow, depended on the helium coolant flow rates and the distance 
behind the area of injection. The results were correlated in terms of the wall cooling parameter and the coolant flow-rate parameter, 
where the nondimensional flow rate was referenced to the cooled area up to the downstream position. For the same coolant flow 
rate, lower surface temperatures are achieved with a porous-wall cooling system. However, since flow-rate requirements decrease 
with increasing allowable surface temperatures, the higher allowable wall temperatures of the solid wall as compared to the struc- 
turally weaker porous wall- sharply reduce the flow -rate requirements of a downstream cooling system. Thus, for certain flight 
conditions it is possible to compensate for the lower efficiency of the downstream or solid-wall cooling system. For example, a 
downstream cooling system using solid walls that must be maintained at 1,800 F would require less coolant for Mach numbers 
up to 5.5 than would a porous-wall cooling system for which the walls must be maintained at temperatures less than or equal to 
9000 F. 
Author 
Leading Edges; Porous Walls; Cooling Systems; Coolants; Air Jets; Solid Surfaces; Wall Temperature; Wings; Cooling 

19980227185 NASA Ames Research Center, Moffett Field, CA USA 

Effects of Sweep Angle on the Boundary-Layer Stability Characteristics of an Untapered Wing at Low Speeds 

Boltz, Frederick W., NASA Ames Research Center, USA; Kenyon, George C, NASA Ames Research Center, USA; Allen, Clyde 

Q., NASA Ames Research Center, USA; Oct. 1960; 80p; In English 

Report No.(s): NASA-TN-D-338; No Copyright; Avail: CASI; A05, Hardcopy; A01, Microfiche 

An investigation was conducted in the Ames 12-Foot Low -Turbulence Pressure Tunnel to determine the effects of sweep on 
the boundary-layer stability characteristics of an untapered variable-sweep wing having an NACA 64(2)A015 section normal to 
the leading edge. Pressure distribution and transition were measured on the wing at low speeds at sweep angles of 0, 10, 20, 30, 
40, and 50 deg. and at angles of attack from -3 to 3 deg. The investigation also included flow -visualization studies on the surface 
at sweep angles from to 50 deg. and total pressure surveys in the boundary layer at a sweep angle of 30 deg. for angles of attack 
from -12 to deg. It was found that sweep caused premature transition on the wing under certain conditions. This effect resulted 
from the formation of vortices in the boundary layer when a critical combination of sweep angle, pressure gradient, and stream 
Reynolds number was attained. A useful parameter in indicating the combined effect of these flow variables on vortex formation 
and on beginning transition is the crossflow Reynolds number. The critical values of crossflow Reynolds number for vortex forma- 
tion found in this investigation range from about 135 to 190 and are in good agreement with those reported in previous investiga- 
tions. The values of crossflow Reynolds number for beginning transitions were found to be between 190 and 260. For each 

62 



condition (i.e., development of vortices and initiation of transition at a given location) the lower values in the specified ranges 
were obtained with a light coating of flow -visualization material on the surface. A method is presented for the rapid computation 
of crossflow Reynolds number on any swept surface for which the pressure distribution is known. From calculations based on 
this method, it was found that the maximum values of crossflow Reynolds number are attained under conditions of a strong pres- 
sure gradient and at a sweep angle of about 50 deg. Due to the primary dependence on pressure gradient, effects of sweep in causing 
premature transition are generally first encountered on the lower surfaces of wings operating at positive angles of attack. 
Author 

Variable Sweep Wings; Sweep Angle; Boundary Layer Stability; Wind Tunnel Tests; Flow Visualization; Cross Flow; Aerody- 
namic Stability 

19980227274 NASA Langley Research Center, Hampton, VA USA 

Laminar Heat-Transfer and Pressure Measurements at a Mach Number of 6 on Sharp and Blent 15 deg Half- Angle Cones 

at Angles of Attack Up to 90 cleg 

Conti, Raul J., NASA Langley Research Center, USA; Oct. 1961; 34p; In English 

Report No.(s): NASA-TN-D-962; L-1624; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Two circular conical configurations having 15 deg half -angles were tested in laminar boundary layer at a Mach number of 
6 and angles of attack up to 90 deg. One cone had a sharp nose and a fineness ratio of 1.87 and the other had a spherically blunted 
nose with a bluntness ratio of 0. 1428 and a fineness ratio of 1 .66. Pressure measurements and schlieren pictures of the flow showed 
that near-conical flow existed up to an angle of attack of approximately 60 deg. At angles of attack above 70 deg high-pressure 
areas were present near the base and the bow shock wave was considerably curved. Comparison of the results with simply applied 
theories showed that on the stagnation line pressures may be predicted by Newtonian theory, and heat transfer by local yawed-cyl- 
inder theory based on the yaw angle of the windward generator and the local radius of the cone. Base effects increased the heat 
transfer in a region extending forward approximately 15 to 30 percent of the windward generator. Circumferential pressure dis- 
tributions were higher than the corresponding Newtonian distribution and a better prediction was obtained by modifying the 
theory to match the pressure at 90 deg from the windward generator to that on the surface of the cone at an angle of attack of 
deg. Circumferential heat-transfer distributions were predicted satisfactorily up to about 60 deg from the stagnation line by using 
Lees' heat-flux distribution based on the Newtonian pressure. The effects of nose bluntness at large angles of attack were very 
small in the region beyond two nose radii from the point of tangency. 
Author 

Laminar Boundary Layer; Heat Transfer; Pressure Measurement; Half Cones; Conical, Flow; Angle of Attack; Pressure Distribu- 
tion; Stagnation Pressure 

19980227276 NASA Langley Research Center, Hampton, VA USA 

Pressure Loads Produced on a Flat-Plate Wing by Rocket Jets Exhausting in a Spanwise Direction Below the Wing and 

Perpendicular to a Free-Stream Flow of Mach Number 2.0 

Falanga, Ralph A., NASA Langley Research Center, USA; Janos, Joseph J., NASA Langley Research Center, USA; May 1961; 

46p; In English 

Report No.(s): NASA-TN-D-893; L-1614; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

An investigation at a Reynolds number per foot of 14.4 x 10(exp 6) was made to determine the pressure loads produced on 
a flat-plate wing by rocket jets exhausting in a spanwise direction beneath the wing and perpendicular to a free-stream flow of 
Mach number 2.0. The ranges of the variables involved were (1) nozzle types - one sonic (jet Mach number of 1.00), two super- 
sonic (jet Mach numbers of 1.74 and 3.04),. and one two-dimensional supersonic (jet Mach number of 1.71); (2) vertical nozzle 
positions beneath the wing of 4, 8 and 12 nozzle-throat diameters; and (3) ratios of rocket-chamber total pressure to free- stream 
static pressure from to 130. The incremental normal force due to jet interference on the wing varied from one to two times the 
rocket thrust and generally decreased as the pressure ratio increased. The chordwise coordinate of the incremental-normal-force 
center of pressure remained upstream of the nozzle center line for the nozzle positions and pressure ratios of the investigation. 
The chordwise coordinate approached zero as the jet vertical distance beneath the wing increased. In the spanwise direction there 
was little change due to varying rocket-jet position and pressure ratio. Some boundary-layer flow separation on the wing was 
observed for the rocket jets close to the wing and at the higher pressure ratios. The magnitude of the chordwise and spanwise pres- 
sure distributions due to jet interference was greatest for rocket jets close to the wing and decreased as the jet was displaced farther 
from the wing. The design procedure for the rockets used is given in the appendix. 
Author 

Aerodynamic Interference; Boundary Layer Separation; Center of Pressure; Pressure Distribution; Pressure Ratio; Supersonic 
Jet Flow; Rocket Thrust; Boundary Layer Flow 

63 



19980227394 NASA Lewis Research Center, Cleveland, OH USA 

Turbulence Studies of a Rectangular Slotted Noise-Suppressor Nozzle 

Laurence, James C, NASA Lewis Research Center, USA; Sep. 1960; 90p; In English 

Report No.(s): NASA-TN-D-294; E-384; No Copyright; Avail: CASI; A05, Hardcopy; A01, Microfiche 

The problem of noise suppression of turbojet engines has shown a need for turbulence data within the flow field of various 
types of nozzles used in ad hoc investigations of the sound power. The result of turbulence studies in a nozzle configuration of 
four parallel rectangular slots is presented in this report with special attention to the effect of the spacing of the nozzles on the 
intensity of turbulence, scale of turbulence, spectrum of turbulence, and the mean stream velocity. Taylor's hypothesis, which 
describes the convection of the turbulence eddies, was tested and found correct within experimental error and certain experimental 
and theoretical limitations. The convection of the pressure patterns was also investigated, and the value of the convection velocity 
was found to be about 0.43 times the central core velocity of the jets. The effect of the spacing-to-width ratio of the nozzles upon 
the turbulence intensity, the scale of turbulence, and the spectral distribution of the noise was found in general to produce a maxi- 
mum change for spacing-to-width ratios of 1.5 to 2.0. These changes may be the cause of the reduction in sound power reported 
for similar full-scale nozzles and test conditions under actual (static) engine operation. A noise reduction parameter is defined 
from Lighthill's theory which gives qualitative agreement with experiments which show the noise reduction is greatest for spac- 
ing-to-width ratios of 1.5 to 2.0. 
Author 
Noise Reduction; Turbulence; Nozzles; Turbojet Engines; Flow Velocity; Full Scale Tests; Flow Distribution 

19980227317 Army Tank-Automotive Research and Development Command, Warren, MI USA 

Lab Test of MIPS Turbodyne II Precleaner with Scavenge Blower Motor, Jan. - May. 1997 

Richard, Michael; McDuffee, Michael; Sierpien, Larry; Margrif, Frank; Jul. 1998; 13 lp; In English 

Report No.(s): AD-A350739; TARDEC-TR- 13752; No Copyright; Avail: CASI; A07, Hardcopy; A02, Microfiche 

An engine induction air precleaner system designed for the MIPS was lab tested at both TARDEC and SWRI to measure pres- 
sure drop, efficiency and particle size determination. The Turbodyne II Precleaner is a two-stage precleaner installed up-stream 
of turbocharged diesel engines. A self -cleaning rotating barrier filter is the second component of the Turbodyne II self-cleaning 
air filter (SCAF) system and is installed after the turbocharger. Two previous tests of the Turbodyne II SCAF were conducted: 
(1) Reference Appendix A, report page and abstract and (2) Reference Appendix B, report page and abstract Test results showed 
in general: (1) Turbocharger degradation does not occur when exposed to precleaned air and (2) some minor difficulty occurred 
in achieving normal efficiency requirements and/or pressure drop limits across SCAF barrier filter for up to 200 hours. TARDEC 
and SwRI pressure drop lab tests were in agreement reaching a maximum of 11.2 to 11.4 inches of water at rated flow of 2600 
cfm. Likewise efficiency testing at TARDEC and SwRI conducted on PTI coarse test dust was nearly in agreement with TARDEC 
obtaining an average overall efficiency of 98.15% compared to the slightly higher average overall efficiency of 98.624% obtained 
by SwRI. Turbodyne II precleaner particle size determination tests conducted by SwRI (See Appendix G, report) showed for three 
dust concentrations (zero visibility, half zero visibility and quarter zero visibility) separation efficiency at low concentration levels 
becomes more sensitive to airflow. For the three dust concentrations tested, test results showed the precleaner had an effective 
cut size ranging from about 3 to 6.5 microns depending on concentration and airflow rate. The cut size is the particle size where 
the probability of collection is 50%. 
DTIC 
Blowers; Cleaning; Air Filters; Superchargers 

19980227345 Physical Research, Inc., Kirkland, WA USA 

Development of an Improved Magneto-Optic/Eddy-Current Imager Final Report 

Thome, David K., Physical Research, Inc., USA; Apr. 1998; 54p; In English 

Contract(s)/Grant(s):DTRS57-95-C-00086 

Report No.(s): AD-A350709; DOT/FAA/AR-97/37; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

Magneto-optic/eddy-current imaging technology has been developed and approved for inspection of cracks in aging aircraft. 
This relatively new nondestructive test method gives the inspector the ability to quickly generate real-time eddy-current images 
of large surface areas. An earlier Phase I Small Business Innovative Research (SBIR) program demonstrated the ability to generate 
improved, complete, real-time magneto-optic/eddy-current images of subsurface corrosion and cracking. Multidirectional eddy- 
current excitation, enhanced low-frequency operation, improved electromagnetic shielding, image processing, and sensor 
improvement were all demonstrated or evaluated. Favorable results from Phase I led to this Phase II SBIR program. The Phase 
II research has resulted in the development of a next generation prototype magneto-optic imager (MOI) with multidirectional 
eddy-current excitation, remotely programmable system settings, on-screen display of system setup information, and improved 

64 



shielding for enhanced images. Some of these new features have already been successfully incorporated into an improved imager, 

the MOI 303, which is now commercially available. 

DTIC 

Eddy Currents; Image Processing; Imaging Techniques; Magneto-Optics; Corrosion; Nondestructive Tests; Aircraft Mainte- 
nance; Images 

19980227349 NASA Lewis Research Center, Cleveland, OH USA 

Estimate of Shock Standoff Distance Ahead of a General Stagnation Point 

Reshotko, Eli, NASA Lewis Research Center, USA; Aug. 1961; 18p; In English 

Report No.(s): NASA-TN-D-1050; E-1278; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The shock standoff distance ahead of a general rounded stagnation point has been estimated under the assumption of a 
constant-density-shock layer. It is found that, with the exception of almost -two-dimensional bodies with very strong shock waves, 
the present theoretical calculations and the experimental data of Zakkay and Visich for toroids are well represented by the relation 
Delta-3D/R(s) = ((Delta-ax sym)/(R(s))/(2/(K+l))) where Delta is the shock standoff distance, R(s),x is the smaller principal 
shock radius, and K is the ratio of the smaller to the larger of the principal shock radii. 
Author 

Shock Waves; Two Dimensional Bodies; Blunt Bodies; Stagnation Point; Inviscid Flow; Aerodynamic Configurations; Symmetri- 
cal Bodies 

19980227352 NASA Langley Research Center, Hampton, VA USA 

A Method of Solution with Tabulated Results For the Attached Oblique Shock -Wave System for Surfaces at Various Angles 

of Attack, Sweep, and Dihedral in an Equilibrium Real Gas Including the Atmosphere 

Trimpi, Robert L., NASA Langley Research Center, USA; Jones, Robert A., NASA Langley Research Center, USA; 1960; 142p; 

In English 

Report No.(s): NASA-TR-R-63; No Copyright; Avail: CASI; A07, Hardcopy; A02, Microfiche 

A new method of solution is derived from basic physical considerations. Results are tabulated for the following ranges: angle 
of attack, deg to 65 deg; angle of sweep, deg to 75 deg; angle of dihedral, deg to 30 deg; Mach number, 3 to 30;and "effective 
specific -heat ratio "parameter, 1.10 to 1.67. Both the method and tabulated solutions are easily adaptable to flight in any gas or in 
the atmosphere of any planet. An illustrative example is presented based on the 1956 ARDC model atmosphere. 
Author 
Angle of Attack; Real Gases; Atmospheric Models; Dihedral Angle 

19980227409 NASA Langley Research Center, Hampton, VA USA 

Configuration Factors for Exchange of Radiant Energy Between Axisymmetrical Sections of Cylinders, Cones, and Hemi- 
spheres and Their Bases 

Buschman, Albert J., Jr., NASA Langley Research Center, USA; Pittman, Claud M., NASA Langley Research Center, USA; Oct. 
1961; 48p; In English 
Report No.(s): NASA-TN-D-944; L-992; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Radiation-interchange configuration factors are derived for axisymmetrical sections of cylinders, cones, and hemispheres 
radiating internally to annular and circular sections of their bases and to other axisymmetrical sections. The general procedure 
of obtaining configuration factors is outlined and the results are presented in the form of equations, tables, and figures. 
Author 
Heat Transfer; Heat Transfer Coefficients; Cones; Aerodynamic Configurations; Radiant Heating; Cylindrical Bodies 

19980227498 NASA Langley Research Center, Hampton, VA USA 

Local Aerodynamic Heat Transfer and Boundary-Layer Transition on Roughened Sphere-Ellipsoid Bodies at Mach 

Number 3.0 

Deveikis, William D., NASA Langley Research Center, USA; Walker, Robert W., NASA Langley Research Center, USA; Aug. 

1961; 26p; In English 

Report No.(s): NASA-TN-D-907; L-1393; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A wind-tunnel investigation was made to determine heat -transfer distributions on three steel sphere-ellipsoid bodies with sur- 
face roughnesses of 5, 100, and 200 microinches. Tests were conducted in the Langley 9- by 6-foot thermal structures tunnel at 
a Mach number of 3.0, free-stream Reynolds numbers (based on model spherical diameter) of 4.25 x 10(exp 6) and 2.76 x 10(exp 
6), and at a stagnation temperature of 650 F. Pressure distributions were obtained also on a fourth model. The results indicated 

65 



that the combination of surface roughness and boundary -layer cooling tended to promote early transition and nullify the advan- 
tages attributable to the blunt shape of the model for reducing local temperatures. Good correlation between experimental heating 
rates and those calculated from laminar theory was achieved up to the start of boundary-layer transition. The correlation also was 
good with the values predicted by turbulent theory for surface stations downstream from the 45 deg. station. 
Author 

Aerodynamic Heat Transfer; Boundary Layer Transition; Spheres; Supersonic Speed; Wind Tunnel Tests; Surface Roughness; 
Free Flow; Cooling 

19980227751 NASA Langley Research Center, Hampton, VA USA 

Free-Flight Investigation of Heat Transfer to an Unswept Cylinder Subjected to an Incident Shock and Flow Interference 

from ao Upstream Body at Mach Numbers up to 5.50 

Carter, Howard S., NASA Langley Research Center, USA; Carr, Robert E., NASA Langley Research Center, USA; Oct. 1961; 

34p; In English 

Report No.(s): NASA-TN-D-988; L-879; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Heat-transfer rates have been measured in free flight along the stagnation line of an unswept cylinder mounted transversely 
on an axial cylinder so that the shock wave from the hemispherical nose of the axial cylinder intersected the bow shock of the 
unswept transverse cylinder. Data were obtained at Mach numbers from 2.53 to 5.50 and at Reynolds numbers based on the trans- 
verse cylinder diameter from 1.00 x 10(exp 6) to 1.87 x 10(exp 6). Shadowgraph pictures made in a wind tunnel showed that the 
flow field was influenced by boundary -layer separation on the axial cylinder and by end effects on the transverse cylinder as well 
as by the intersecting shocks. Under these conditions, the measured heat-transfer rates had inconsistent variations both in magni- 
tude and distribution which precluded separating the effects of these disturbances. The general magnitude of the measured heating 
rates at Mach numbers up to 3 was from 0.1 to 0.5 of the theoretical laminar heating rates along the stagnation line for an infinite 
unswept cylinder in undisturbed flow. At Mach numbers above 4 the measured heating rates were from 1.5 to 2 times the theoreti- 
cal rates. 
Author 

Free Flight; Wind. Tunnel Tests; Supersonic Speed; Shock Waves; Heat Transfer; Aerodynamic Heating; Bow Waves; Cylindrical- 
Bodies; Unswept Wings 

19980227735 NASA Langley Research Center, Hampton, VA USA 



Preliminary Investigation of an Underwater Ramjet Powered by Compressed Air 

Mottard, Elmo J., NASA Langley Research Center, USA; Shoemaker, Charles J., NASA Langley Research Center, USA; Dec. 

1961; 36p; In English 

Report No.(s): NASA-TN-D-991; L-1249; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Part I contains the results of a preliminary experimental investigation of a particular design of an underwater ramjet or hydro- 
duct powered by compressed air. The hydroduct is a propulsion device in which the energy of an expanding gas imparts additional 
momentum to a stream of water through mixing. The hydroduct model had a fineness ratio of 5.9, a maximum diameter of 3.2 
inches, and a ratio of inlet area to frontal area of 0.32. The model was towed at a depth of 1 inch at forward speeds between 20 
and 60 feet per second for airflow rates from 0. 1 to 0.3 pound per second. Longitudinal force and pressures at the inlet and in the 
mixing chamber were determined. The hydroduct produced a positive thrust-minus-drag force at every test speed. The force and 
pressure coefficients were functions primarily of the ratio of weight airflow to free-stream velocity. The maximum propulsive 
efficiency based on the net internal thrust and an isothermal expansion of the air was approximately 53 percent at a thrust coeffi- 
cient of 0. 10. The performance of the test model may have been influenced by choking of the exit flow. Part II is a theoretical 
development of an underwater ramjet using air as "fuel." The basic assumption of the theoretical analysis is that a mixture of water 
and air can be treated as a compressible gas. More information on the properties of air-water mixtures is required to confirm this 
assumption or to suggest another approach. A method is suggested from which a more complete theoretical development, with 
the effects of choking included, may be obtained. An exploratory computation, in which this suggested method was used, indicated 
that the effect of choked flow on the thrust coefficient was minor. 
Author 
Ramjet Engines; Propulsion; Compressed Air; Free Flow; Propulsive Efficiency; Thrust 

19980227763 NASA Langley Research Center, Hampton, VA USA 

Cross-Sectional Deformations of Monocoque Beams and Their Effects on the Natural Vibration Frequencies 
Thomson, Robert G., NASA Langley Research Center, USA; Kruszewski, Edwin T., NASA Langley Research Center, USA; Dec. 
1961; 50p; In English 

66 



Report No.(s): NASA-TN-D-987; L-1444; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

The variational principle, differential equations, and boundary conditions governing the cross-sectional distortions due to 
inertia loading of a two-dimensional model of a thin monocoque wing are shown. A theoretical analysis of this simplified model 
is made in order to determine the nature of the coupling between the cross-sectional modes and the spanwise deformation modes. 
General solutions are obtained in finite-difference form for arbitrary cross sections and an exact solution is presented for a parabol- 
ic-arc cross section of constant cover thickness. The application of these results in evaluating the coupled frequencies of the actual 
structure is discussed. Frequencies evaluated for a parabolic-arc monocoque beam show good agreement with experimental val- 
ues. 
Author 

Monocoque Structures; Aircraft Structures; Dynamic Structural Analysis; Finite Difference Theory; Structural Vibration; 
Deformation 

19980227795 NASA Langley Research Center, Hampton, VA USA 

Experimental Investigation at Mach Number 3.0 of the Effects of Thermal Stress and Buckling on the Flutter of Four-Bay 

Aluminum Alloy Panels with Length-Width Ratios of 10 

Dixon, Sidney C, NASA Langley Research Center, USA; Griffith, George E., NASA Langley Research Center, USA; Bohon, 

Herman L., NASA Langley Research Center, USA; Oct. 1961; 30p; In English 

Report No.(s): NASA-TN-D-921; L-1265; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Skin-stiffener aluminum alloy panels consisting of four bays, each bay having a length-width ratio of 10, were tested at a Mach 
number of 3.0 at dynamic pressures ranging from 1,500 psf to 5,000 psf and at stagnation temperatures from 300 F to 655 F. The 
panels were restrained by the supporting structure in such a manner that partial thermal expansion of the skins could occur in both 
the longitudinal and lateral directions. A boundary faired through the experimental flutter points consisted of a flat-panel portion, 
a buckled-panel portion, and a transition point at the intersection of the two boundaries. In the region where a panel must be flat 
when flutter occurs, an increase in panel skin temperature (or midplane compressive stress) makes the panel more susceptible to 
flutter. In the region where a panel must be buckled when flutter occurs, the flutter trend is reversed. This reversal in trend is attrib- 
uted to the panel postbuckling behavior. 
Author 
Aluminum, Alloys; Panels; Supersonic Speed; Buckling; Thermal Stresses; Panel Flutter; Vibrational Stress; Aeroelasticity 

14 
LIFE SCIENCES 

Includes life sciences (general); aerospace medicine; behavioral sciences; man/system technology and life support; and space biology. 

19981)223621 NASA Ames Research Center, Moffett Field, CA USA 

Centrifuge Study of Pilot Tolerance to Acceleration and the Effects of Acceleration on Pilot Performance 

Creer, Brent Y., NASA Ames Research Center, USA; Smedal, Harald A., NASA Ames Research Center, USA; Wingrove, Rodney 

C, NASA Ames Research Center, USA; Nov. 1960; 38p; In English 

Report No.(s): NASA-TN-D-337; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

A research program the general objective of which was to measure the effects of various sustained accelerations on the control 
performance of pilots, was carried out on the Aviation Medical Acceleration Laboratory centrifuge, U.S. Naval Air Development 
Center, Johnsville, PA. The experimental setup consisted of a flight simulator with the centrifuge in the control loop. The pilot 
performed his control tasks while being subjected to acceleration fields such as might be encountered by a forward-facing pilot 
flying an atmosphere entry vehicle. The study was divided into three phases. In one phase of the program, the pilots were subjected 
to a variety of sustained linear acceleration forces while controlling vehicles with several different sets of longitudinal dynamics. 
Here, a randomly moving target was displayed to the pilot on a cathode-ray tube. For each combination of acceleration field and 
vehicle dynamics, pilot tracking accuracy was measured and pilot opinion of the stability and control characteristics was recorded. 
Thus, information was obtained on the combined effects of complexity of control task and magnitude and direction of acceleration 
forces on pilot performance. These tests showed that the pilot's tracking performance deteriorated markedly at accelerations 
greater than about 4g when controlling a lightly damped vehicle. The tentative conclusion was also reached that regardless of the 
airframe dynamics involved, the pilot feels that in order to have the same level of control over the vehicle, an increase in the vehicle 
dynamic stability was required with increases in the magnitudes of the acceleration impressed upon the pilot. In another phase, 
boundaries of human tolerance of acceleration were established for acceleration fields such as might be encountered by a pilot 
flying an orbital vehicle. A special pilot restraint system was developed to increase human tolerance to longitudinal decelerations. 

67 



The results of the tests showed that human tolerance of longitudinal deceleration forces was considerably improved through use 

of the special restraint system. 

Author 

Pilot Performance; Human Tolerances; Flight Simulators; Deceleration; Dynamic Stability; Atmospheric Entry; Centrifuges 

19980223933 Naval Aerospace Medical Research Lab., Pensacola, FL USA 

Calculating A Helicopter Pilot's Instrument Scan Pattern from Discrete, 60-Hz Measures of the Line-of-Sighl: The Evalu- 
ation of an Algorithm 
Jun. 17, 1998; 33p; In English 
Report No.(s): AD-A350657; NAMRL-1403; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

In order obtain data to develop and evaluate theories relating instrument scanning to flight performance we recorded the line 
of sight (LOS) of student naval helicopter pilots as they flew prescribed maneuvers in a motion-based, high fidelity, instrument 
training simulator. These LOS data were discrete, 60 Hz samples of eye pointing. For some types of analysis it is helpful to think 
of a scan pattern as a sequence of fixations and to use an averaging algorithm to transform the 60 Hz data into such a sequence, 
a scan path. An appropriate algorithm was identified, developed and evaluated. As part of this evaluation, we developed a String 
Similarity measure, SS, a measure of the similarity between two scan paths. The evaluation of the algorithm, consisting of observ- 
ing the algorithm's output as a function of the algorithm's parameter values, showed that the algorithm behaved in a sensible fash- 
ion, logically consistent with the input data. This increased our confidence in our implementation of the fixation algorithm. The 
SS metric proved to be an informative, useful tool that may have addition uses in the analysis scanning behavior and flight perfor- 
mance. 
DTIC 
Helicopters; Scanners; Flight Instruments; Line of Sight; Algorithms; Aircraft Pilots 

19980223984 Washington Univ., Seattle, WA USA 

The Adaptive Effects of Virtual Interfaces: Vestibulo-Ocular Reflex arid Simulator Sickness 

Draper, Mark H., Washington Univ., USA; Aug. 07, 1998; 345p; In English 

Report No.(s): AD-A350767; AFIT-98-021D; No Copyright; Avail: CASI; A15, Hardcopy; A03, Microfiche 

Current virtual interfaces imperfectly simulate the motion dynamics of the real world. Conflicting visual and vestibular cues 
of self-motion are believed to result in vestibulo-ocular reflex (VOR) adaptations and simulator sickness, which raises health and 
safety issues surrounding virtual environment (VE) exposure. Four experiments were conducted to examine the effects of conflict- 
ing visual-vestibular cues through employment of typically occurring virtual interface scenarios. Subjects were exposed for 30 
minutes to a head-coupled virtual interface, completing visual search tasks using active, unrestricted head movement rotations. 
DTIC 
Virtual Reality ; Visual Perception; Motion Sickness; Flight Simulators; Reflexes 

19980227270 Army Aeromedical Research Lab., Fort Rucker, AL USA 

Effects of Head-Supported Devices on Female Aviators dnring Simulated Helicopter Missions Annual Report 

Alem, Nabih, Army Aeromedical Research Lab., USA; May 1998; 122p; In English 

Report No.(s): AD-A350472; No Copyright; Avail: CASI; A06, Hardcopy; A02, Microfiche 

This report describes the work completed during the first project year of this research study. The objective of the study is to 
identify safe weight and location limits of head-supported devices worn by female aviators during simulated helicopter rides. The 
working hypothesis is that female pilots will tolerate some range of HSD weight moments beyond which their biomechanical and 
performance responses will deteriorate. The report contains a review of relevant studies followed by detailed description of the 
experimental and analytical procedures. 
DTIC 
Aircraft Pilots; Females; Helicopters; Helmets 

19980227326 Air Force Research Lab., Human Effectiveness Directorate, Wright-Patterson AFB, OH USA 

Building the LeM2*R3 Model of Pilot Trust and Dynamic Workload Allocation; .4 Transition of Theory and Empirical 

Observations to Cockpit Demonstration Final Report, Jan. 1994 - Oct. 1997 

Raeth, Peter G.; Reising, John M.; Feb. 1998; 102p; In English 

Contract(s)/Grant(s): Proj-2403 

Report No.(s): AD-A350481; AFRL-HE-WP-TR-1998-0046; No Copyright; Avail: CASI; A06, Hardcopy; A02, Microfiche 

68 



For pilots to accept active decision aids during complex flight scenarios, it is essential that the automation work is in synergy 
with aircrew. To accomplish this, the automation must go well beyond menu and macro selections, where the pilot must explicitly 
tell the automation what to do and when to do it. It must also transcend "mother may I" approaches, where the automation asks 
for permission to proceed, to these traditional barriers, the automation needs a sense of how the pilot will react in a given situation 
and, based on that reaction, how much of the workload could be allocated to the automation at any given time. For this purpose, 
the authors reviewed the literature on human factors and dynamic function allocation. This literature provided a wealth of informa- 
tion on this topic. Based on the current state of the art in this topic area, the authors developed and tested a dynamic model of pilot 
trust and workload allocation. This "full degrees of freedom" model transitions human factors theory, as it exists today, into an 
engineering application. The resulting model can be combined with other information obtained from static and continuous pro- 
cesses to divide the workload and minimize cognitive overload. 
DTIC 
Cockpits; Human Factors Engineering; Decision Support Systems; Artificial Intelligence; Decision Making; Flight Simulation 



15 
MATHEMATICAL AND COMPUTER SCIENCES 

Includes mathematical and computer sciences (general); computer operations and hardware; computer programming and software; 
computer systems; cybernetics; numerical analysis; statistics and probability; systems analysis; and theoretical mathematics. 

19980227167 Army Command and General Staff Coll., Fort Leavenworth, KS USA 

F-16 Peacetime Training for Combat Operations 

Roosa, John D., Army Command and General Staff Coll., USA; Jun. 05, 1998; 98p; In English 

Report No.(s): AD-A350054; No Copyright; Avail: CASI; A05, Hardcopy; A02, Microfiche 

This study investigates the relationship between peacetime F-16 training and expected combat operations. The F-16 is the 
primary interdiction platform in the USA Air Force. F-16 pilots fly peacetime training sorties to maintain proficiency, develop 
tactics and complete evaluations. The training activities accomplished on these missions are designed to prepare the pilot for suc- 
cessful combat employment. A training program ensures each pilot completes the necessary amount of sorties and events to 
achieve combat ready status. This study analyzes the components of the peacetime training program and their overall applicability 
for future conflict. The study encompasses the entire training program from higher headquarters directives down to specific flying 
sorties. 
DTIC 
Peacetime; Flight Training; F-16 Aircraft; Education 

19980227173 Army Command and General Staff Coll., Fort Leavenworth, KS USA 
Premobilization Proficiency of USA Army Reserve Attack Helicopter Battalions 
Gruenwald, David L., Army Command and General Staff Coll., USA; Jun. 05, 1998; 85p; In English 
Report No.(s): AD-A349995; No Copyright; Avail: CASI; A05, Hardcopy; A01, Microfiche 

Attack helicopter battalions are combat maneuver units that conduct supporting attacks which aid, protect, and compliment 
other maneuver forces by destroying massed enemy mechanized forces and other enemy forces with aerial firepower, mobility, 
and shock effect. They are employed as a battalion in order to provide the commander with this highly mobile and lethal destruc- 
tion capability. The fundamentals of attack helicopter operations do not change by component. Reserve Component attack heli- 
copter battalions are expected to perform attack helicopter operations to the same level of proficiency or standard as the Active 
Component. Currently, there is conflicting guidance published by Forces Command as to what level of proficiency aviation units 
in the Reserve component should train to in premobilization in order to prepare for their wartime mission. This study examines 
the ability of USA Army Reserve (US AR) attack helicopter units to maintain proficiency at the battalion level in a premobilization 
environment. It focuses on the resources available to Reserve units and the training requirements placed on a unit. It concludes 
with an analysis of a US AR attack helicopter unit's ability to execute all training requirements in the time available to them each 
training year. It offers recommendations on possible alternative training strategies and provides suggestions for further research. 
DTIC 
Helicopters; Armed Forces (USA); Attack Aircraft; Education 

69 



16 
PHYSICS 

Includes physics (general); acoustics; atomic and molecular physics; nuclear and high-energy; optics; plasma physics; solid-state phys- 
ics; and thermodynamics and statistical physics. 

19980223604 NASA Lewis Research Center, Cleveland, OH USA 

Similarity of Near Noise Fields of Subsonic Jets 

Howes, Walton L., NASA Lewis Research Center, USA; 1961; 56p; In English 

Report No.(s): NASA-TR-R-94; No Copyright; Avail: CASI; A04, Hardcopy; A01, Microfiche 

Similarity relations for frequency pass band, as well as overall, time average pressure fluctuations outside a jet are derived 
and tested using experimental data. Similarity of the pressure fields was found for different jet velocities. Nozzle contour dissimi- 
larity and differing jet temperatures were found to limit seriously the application of the similarity relations, especially near the 
jet nozzle. 
Author 

Subsonic Flow; Air Jets; Engine Noise; Near Fields; Pressure Oscillations; Aerodynamic Noise; Pressure Gradients; Sound 
Pressure 

18 
SPACE SCIENCES 

Includes space sciences (general); astronomy; astrophysics; lunar and planetary exploration; solar physics; and space radation. 

19981)223611 NASA Marshall Space Flight Center, Huntsville, AL USA 

Aerodynamic Analysis of Tektites and Their Parent: Bodies 

Adams, E. W., NASA Marshall Space Flight Center, USA; Huffaker, R. M., NASA Marshall Space Flight Center, USA; 1962; 

48p; In English 

Report No.(s): NASA-TR-R-149; No Copyright; Avail: CASI; A03, Hardcopy; A01, Microfiche 

Experiment and analysis indicate that the button-type australites were derived from glassy spheres which entered or re-en- 
tered the atmosphere as cold solid bodies; in case of average-size specimens, the entry direction was nearly horizontal and the entry 
speed between 6.5 and 11.2 km/sec. Terrestrial origin of such spheres is impossible because of extremely high deceleration rates 
at low altitudes. The limited extension of the strewn fields rules out extraterrestrial origin of clusters of such spheres because of 
stability considerations for clusters in space. However, tektites may have been released as liquid droplets from glassy parent bodies 
ablating in the atmosphere of the earth. The australites then have skipped together with the parent body in order to re-enter as cold 
spheres. Terrestrial origin of a parent body would require an extremely violent natural event. Ablation analysis shows that fusion 
of opaque siliceous stone into glass by aerodynamic heating is impossible. 
Author 
Design Analysis; Tektites; Aerodynamic Heating; Ablation; Aerodynamic Characteristics; Australites 



70 



Subject Terms Index 



ABLATION, 55, 57, 70 

ACCELERATION (PHYSICS), 11 

ACTIVE CONTROL, 51 

ADAPTIVE CONTROL, 43 

AERODYNAMIC BALANCE, 6, 8, 49 

AERODYNAMIC CHARACTER- 
ISTICS, 6, 7, 8, 9, 10, 11, 12, 13, 14, 
15, 16, 17, 19, 25, 27, 28, 30, 31, 32, 
33, 34, 37, 40, 52, 54, 55, 56, 62, 70 

AERODYNAMIC COEFFICIENTS, 3, 7, 
8, 9, 10, 13, 14, 16, 17, 18, 28, 30, 

32, 43, 44, 56 
AERODYNAMIC CONFIGURATIONS, 

7, 8, 11, 16, 18, 33, 45, 49, 56, 60, 65 
AERODYNAMIC DRAG, 2, 4, 5, 7, 15, 

28,30,31,54,61 
AERODYNAMIC FORCES, 14, 16, 56, 

61 
AERODYNAMIC HEAT TRANSFER, 

53, 57, 59, 66 
AERODYNAMIC HEATING, 16, 53, 55, 

56,57,61,66,70 
AERODYNAMIC INTERFERENCE, 13, 

17,63 
AERODYNAMIC LOADS, 9, 18 
AERODYNAMIC NOISE, 6, 59, 70 
AERODYNAMIC STABILITY, 27, 30, 

33, 35, 41, 43, 49, 50, 54, 63 
AERODYNAMIC STALLING, 3, 25 
AERODYNAMICS, 1, 5, 42 
AEROELASTICITY, 17, 42, 47, 67 
AEROMANEUVERING, 14 
AERONAUTICAL ENGINEERING, 1 
AEROSOLS, 24 

AEROSPACE VEHICLES, 14, 56 
AEROSPIKE ENGINES, 40 
AEROTHERMODYNAMICS, 54, 57, 59 
AFTERBODIES, 2, 13, 31, 54, 57 
AH-64 HELICOPTER, 21, 29 
AILERONS, 15 
AIR CUSHION LANDING SYSTEMS, 

31 
AIR FILTERS, 64 
AIRFLOW, 11 
AIR JETS, 62, 70 
AIR NAVIGATION, 20 
AIR TRAFFIC, 19, 22 
AIR TRAFFIC CONTROL, 19, 21, 22, 

23,24 

AIR TRAFFIC CONTROLLERS (PER- 
SONNEL), 23, 24 



AIR TRANSPORTATION, 53 
AIRCRAFT ACCIDENTS, 20 
AIRCRAFT COMPARTMENTS, 33 
AIRCRAFT CONFIGURATIONS, 10, 

30, 36, 37, 49, 50, 55 
AIRCRAFT CONTROL, 16, 27, 34, 36, 

42, 44, 45, 49 
AIRCRAFT DESIGN, 1, 5, 6, 29, 35, 37, 

46 
AIRCRAFT FUELS, 21 
AIRCRAFT INSTRUMENTS, 38, 39 
AIRCRAFT LANDING, 23 
AIRCRAFT MAINTENANCE, 1, 65 
AIRCRAFT MODELS, 9, 19, 43, 45 
AIRCRAFT PERFORMANCE, 4, 23, 35 
AIRCRAFT PILOTS, 38, 68 
AIRCRAFT SAFETY, 20 
AIRCRAFT SPIN, 3, 16, 18, 25 
AIRCRAFT STABILITY, 3, 4, 25, 33, 42, 

44,50 
AIRCRAFT STRUCTURES, 1, 6, 27, 37, 

47, 60, 67 
AIRFOIL OSCILLATIONS, 17 
AIRFOIL PROFILES, 3, 4, 5, 8, 9 
AIRFOILS, 5, 12, 19 
AIRFRAMES, 33 
AIRLINE OPERATIONS, 29 
AIRPORT SECURITY, 20 
AIRPORTS, 19, 20 
AIRSPACE, 22, 23 
ALERTNESS, 23 
ALGORITHMS, 58, 68 
ALL-WEATHER AIR NAVIGATION, 21 
ALL-WEATHER LANDING SYSTEMS, 

21 
ALLOCATIONS, 21 
ALTERNATING CURRENT, 28 
ALTIMETERS, 38, 39 
ALTIMETRY, 38, 39 
ALTITUDE, 20 

ALTITUDE SIMULATION, 50 
ALUMINUM, 24 
ALUMINUM ALLOYS, 67 
ANGLE OF ATTACK, 5, 13, 27, 31, 33, 

38, 56, 63, 65 
APOLLO SPACECRAFT, 57 
APPROACH, 31 
APPROXIMATION, 14 
ARMED FORCES (UNITED STATES), 

69 
ARROW WINGS, 5, 7, 9, 12, 43 
ARTIFICIAL INTELLIGENCE, 47, 69 



ASPECT RATIO, 48 
ATMOSPHERIC ENTRY, 14, 55, 56, 68 
ATMOSPHERIC MODELS, 65 
ATMOSPHERIC PHYSICS, 17 
ATMOSPHERIC TURBULENCE, 42 
ATTACK AIRCRAFT, 69 
AUSTRALITES, 70 
AUTOMATIC CONTROL, 46, 53 
AUTOMATIC PILOTS, 47 
AVIONICS, 38 

AXIAL FLOW TURBINES, 40 
AXISYMMETRIC BODIES, 60 
AXISYMMETRIC FLOW, 14, 60 



B 

B-70 AIRCRAFT, 2 

BACKWARD FACING STEPS, 60 

BALLOONS, 18 

BASE PRESSURE, 5, 13 

BEACONS, 22 

BIBLIOGRAPHIES, 1 

BLADE TIPS, 40 

BLOWERS, 64 

BLOWING, 11, 61 

BLUNT BODIES, 2, 3, 10, 65 

BLUNT LEADING EDGES, 15 

BOATTAILS, 31 

BODIES OF REVOLUTION, 2, 7 

BODY- WING AND TAIL CONFIGU- 
RATIONS, 6, 9, 50 

BODY- WING CONFIGURATIONS, 4, 5, 

6, 8, 10, 12, 14, 37, 43 

BOEING 747 AIRCRAFT, 47 
BOEING 757 AIRCRAFT, 21 
BOMBER AIRCRAFT, 46 
BOMBS (ORDNANCE), 17 
BOOSTER ROCKET ENGINES, 10 
BOOSTGLIDE VEHICLES, 34 
BOUNDARY LAYER CONTROL, 9, 11, 

28, 49, 61 
BOUNDARY LAYER FLOW, 60, 63 
BOUNDARY LAYER SEPARATION, 5, 

11,30,63 
BOUNDARY LAYER STABILITY, 63 
BOUNDARY LAYER TRANSITION, 2, 

15,66 
BOUNDARY LAYERS, 60 
BOW WAVES, 5, 6, 15, 66 
BUCKLING, 67 



ST-1 



C 123 AIRCRAFT, 59 

C 135 AIRCRAFT, 28 

CAMBERED WINGS, 4, 6, 7, 8, 9 

CANARD CONFIGURATIONS, 27, 44, 
47 

CARGO, 24 

CAVITATION FLOW, 7 

CENTER OF GRAVITY, 33 

CENTER OF PRESSURE, 13, 63 

CENTRIFUGES, 46, 68 

CIRCULAR CONES, 60, 61 

CIRCULAR CYLINDERS, 60 

CIRCULAR ORBITS, 54 

CIVIL AVIATION, 20 

CLEANING, 64 

CLEARANCES, 40 

CLIMBING FLIGHT, 2 

COCKPIT SIMULATORS, 46 

COCKPITS, 69 

COEFFICIENT OF FRICTION, 59 

COLLISIONS, 35 

COMBUSTION PRODUCTS, 58 

COMMERCIAL AIRCRAFT, 24, 29 

COMPRESSED AIR, 66 

COMPRESSORS, 52 

COMPUTATIONAL FLUID DYNAM- 
ICS, 17, 40 

COMPUTER AIDED DESIGN, 44 

COMPUTER PROGRAMS, 35 

COMPUTER SYSTEMS PROGRAMS, 
21 

CONDUCTIVE HEAT TRANSFER, 16 

CONES, 10, 16, 65 

CONFERENCES, 17 

CONICAL BODIES, 18 

CONICAL CAMBER, 9 

CONICAL FLOW, 63 

CONICAL SHELLS, 16 

CONICS, 54 

CONSTRICTIONS, 20 

CONTAINMENT, 39, 41 

CONTROL EQUIPMENT, 43 

CONTROL STABILITY, 3, 16, 42 

CONTROL SURFACES, 27, 43, 48 

CONTROL SYSTEMS DESIGN, 44, 52, 
53 

CONTROL THEORY, 43, 44 

CONTROLLABILITY, 2, 42, 44 

CONTROLLERS, 22, 24, 43, 51 

COOLANTS, 62 

COOLING, 11, 62, 66 

COOLING SYSTEMS, 62 

CORROSION, 65 

COST ANALYSIS, 29 



COUNTER ROTATION, 59 
CRASHES, 21, 33 
CRITICAL VELOCITY, 48 
CROSS FLOW, 62, 63 
CRUCIFORM WINGS, 15 
CUES, 46 
CUSHIONS, 31 

CYLINDRICAL BODIES, 2, 10, 11, 15, 
17,31,65,66 



DAMAGE ASSESSMENT, 33 

DAMPING, 48 

DATA ACQUISITION, 26 

DATA LINKS, 21, 22, 28 

DATA STORAGE, 26 

DECELERATION, 14, 68 

DECISION MAKING, 69 

DECISION SUPPORT SYSTEMS, 69 

DEFLECTION, 13, 19, 48 

DEFORMATION, 33, 67 

DELTA WINGS, 11, 12, 13, 25, 27, 30, 

34,41,44,50,51 
DESCENT, 31 

DESCENT TRAIECTORIES, 25, 53, 56 
DESIGN ANALYSIS, 27, 30, 41, 47, 53, 

54, 55, 59, 70 
DETECTION, 5 

DIFFERENTIAL EQUATIONS, 56 
DIHEDRAL ANGLE, 65 
DIRECTIONAL CONTROL, 43, 44 
DIRECTIONAL STABILITY, 28, 33, 34, 

44, 45, 50 
DISPLAY DEVICES, 51 
DISTORTION, 32 
DRAG, 2, 14, 57 
DRAG REDUCTION, 16, 54, 61 
DRIFT (INSTRUMENTATION), 38 
DRONE VEHICLES, 27 
DROP TESTS, 33, 35 
DUCTED FANS, 36, 49 
DYNAMIC CHARACTERISTICS, 2, 27, 

53 
DYNAMIC CONTROL, 46 
DYNAMIC PRESSURE, 48 
DYNAMIC STABILITY, 34, 35, 50, 68 
DYNAMIC STRUCTURAL ANALYSIS, 

67 
DYNAMIC TESTS, 61 



EDDY CURRENTS, 65 
EDUCATION, 69 



ELECTRIC MOTORS, 59 

END PLATES, 61 

ENGINE DESIGN, 41 

ENGINE NOISE, 70 

ENGINE PARTS, 21,41 

EQUATIONS OF MOTION, 55 

ERROR ANALYSIS, 22 

ERRORS, 5, 22, 32 

ESTIMATING, 58 

EVALUATION, 51 

EXHAUST NOZZLES, 5 

EXTERNALLY BLOWN FLAPS, 9, 11, 

28 
EXTINGUISHING, 21 



F-16 AIRCRAFT, 69 

F-5 AIRCRAFT, 16 

FABRICS, 19 

FAIRINGS, 7 

FAN BLADES, 59 

FEEDBACK, 27 

FEEDBACK CONTROL, 44 

FEMALES, 68 

FIGHTER AIRCRAFT, 3, 28, 29, 35, 46 

FINENESS RATIO, 61 

FINITE DIFFERENCE THEORY, 67 

FINS, 18,46 

FIRE EXTINGUISHERS, 21 

FIRE PREVENTION, 21, 24 

FIXED WINGS, 38 

FLAPPING, 37 

FLAPS (CONTROL SURFACES), 37 

FLAT PLATES, 16, 17, 59 

FLIGHT CHARACTERISTICS, 4, 24, 

25, 34, 36, 49, 57 
FLIGHT CONDITIONS, 48 
FLIGHT CONTROL, 43, 44, 46 
FLIGHT INSTRUMENTS, 38, 68 
FLIGHT MANAGEMENT SYSTEMS, 

19 
FLIGHT PATHS, 37, 48 
FLIGHT SAFETY, 20 
FLIGHT SIMULATION, 69 
FLIGHT SIMULATORS, 44, 46, 68 
FLIGHT TESTS, 23, 25, 26, 27, 28, 30, 

32, 33, 34, 36, 38, 39, 40, 46, 55, 57 
FLIGHT TRAINING, 69 
FLOW CHARACTERISTICS, 59, 62 
FLOW DISTRIBUTION, 15, 17, 19, 64 
FLOW GEOMETRY, 62 
FLOW VELOCITY, 17, 64 
FLOW VISUALIZATION, 2, 63 
FLUID FLOW, 59 



ST-2 



FLUTTER ANALYSIS, 18, 42, 46, 47, 

48 
FORCE DISTRIBUTION, 52 
FOREBODIES, 16, 50 
FRAGMENTS, 39, 41 
FREE FLIGHT, 27, 33, 34, 66 
FREE FLOW, 5, 15, 18, 60, 66 
FREQUENCY RESPONSE, 32 
FRICTION DRAG, 16 
FRUSTUMS, 10 

FULL SCALE TESTS, 58, 61, 64 
FUNCTIONAL DESIGN SPECIFI- 
CATIONS, 20 
FUSELAGES, 7, 33, 50 



GAS BAGS, 25 
GAS INJECTION, 16 
GAS TURBINES, 40 
GENETIC ALGORITHMS, 44 
GLIDE LANDINGS, 56 
GLIDERS, 33, 36 

GLOBAL POSITIONING SYSTEM, 23 
GRAPHICAL USER INTERFACE, 35 
GRAPHS (CHARTS), 20 
GROUND EFFECT MACHINES, 31 
GUIDANCE (MOTION), 47 
GUSTS, 42 



H 

HALF CONES, 16, 63 

HANGARS, 58 

HEAT SHIELDING, 49 

HEAT TRANSFER, 15, 40, 57, 59, 63, 

65,66 
HEAT TRANSFER COEFFICIENTS, 57, 

65 
HELICOPTER ENGINES, 41 
HELICOPTERS, 7, 29, 35, 36, 48, 68, 69 
HELIUM, 16 
HELMETS, 68 

HEMISPHERICAL SHELLS, 16 
HIGH SPEED, 36 
HIGH TEMPERATURE TESTS, 61 
HORIZONTAL FLIGHT, 34 
HORIZONTAL TAIL SURFACES, 10, 

46 
HOVERING, 30, 34 
HUBS, 7 
HUMAN FACTORS ENGINEERING, 

20,69 
HUMAN TOLERANCES, 68 
HYDROCHLORIC ACID, 24 



HYDROGEN, 58 
HYPERSONIC GLIDERS, 10, 32 
HYPERSONIC REENTRY, 57 
HYPERSONIC SPEED, 2, 17, 32, 57 
HYPERSONICS, 2, 52, 58 

I 

IDEAL GAS, 18 

IMAGE PROCESSING, 65 

IMAGES, 65 

IMAGING TECHNIQUES, 65 

IMPACT LOADS, 29 

IMPACT TESTS, 33 

IMPELLERS, 59 

INCONEL (TRADEMARK), 61 

INDENTATION, 4 

INDEXES (DOCUMENTATION), 1 

INDICATING INSTRUMENTS, 38 

INFLATABLE STRUCTURES, 19 

INFORMATION TRANSFER, 28, 51 

INFRARED DETECTORS, 28 

INSPECTION, 1 

INSTALLING, 39 

INSTRUMENT FLIGHT RULES, 21 

INSULATED STRUCTURES, 61 

INTERCEPTORS, 35 

INTERFERENCE DRAG, 10 

INTERNAL COMPRESSION INLETS, 

60 
INTERNATIONAL SPACE STATION, 

54 
INVESTIGATION, 29 
INVISCID FLOW, 65 



JET AIRCRAFT, 11,28 
JET AIRCRAFT NOISE, 6 
JET EXHAUST, 5, 13 
JET FLOW, 13, 19 
JP-4 JET FUEL, 54 



LAMINAR BOUNDARY LAYER, 32, 

61,63 
LAMINAR FLOW, 59 
LANDING, 31 

LANDING INSTRUMENTS, 39 
LANDING LOADS, 29 
LATERAL CONTROL, 8, 35, 44, 50 
LATERAL STABILITY, 49, 50, 51 
LAUNCH VEHICLES, 25 



LEADING EDGE SLATS, 11 

LEADING EDGES, 3, 17, 30, 61, 62 

LIFT, 2, 5, 10, 11,28 

LIFT AUGMENTATION, 19 

LIFT DRAG RATIO, 4, 6, 15, 30, 31 

LIFT FANS, 49 

LIFTING REENTRY VEHICLES, 37, 50 

LIFTING ROTORS, 17 

LINE OF SIGHT, 68 

LOADS (FORCES), 12 

LOGISTICS MANAGEMENT, 29 

LONGITUDINAL CONTROL, 37, 46 

LONGITUDINAL STABILITY, 4, 8, 28, 

37,41,43,45,49,61 
LOW ASPECT RATIO, 14, 25, 35 
LOW REYNOLDS NUMBER, 41 
LUNAR EXPLORATION, 50 



M 

MACH NUMBER, 2, 4, 46, 59 

MAGNETO-OPTICS, 65 

MANEUVERABILITY, 10, 35 

MANEUVERS, 3, 47 

MANNED REENTRY, 50, 56, 57 

MANNED SPACECRAFT, 37 

MANUFACTURING, 53 

MARS SURFACE, 54 

MATHEMATICAL MODELS, 47 

MEMORY, 23 

METEOROLOGY, 54 

METHOD OF CHARACTERISTICS, 2 

MISSILE CONTROL, 47 

MISSILE TESTS, 8 

MISSILE TRAJECTORIES, 8 

MISSILES, 47 

MISSION PLANNING, 54 

MODELS, 53 

MOMENT DISTRIBUTION, 16 

MOMENTS, 61 

MONOCOQUE STRUCTURES, 67 

MOTION SICKNESS, 68 



N 

NACELLES, 43 

NATIONAL AIRSPACE SYSTEM, 20 

NAVIGATION INSTRUMENTS, 21 

NAVIGATORS, 31 

NEAR FIELDS, 70 

NEURAL NETS, 47 

NEWTON THEORY, 61 

NEWTONIAN FLUIDS, 61 

NOISE MEASUREMENT, 19 



ST-3 



NOISE REDUCTION, 64 
NONDESTRUCTIVE TESTS, 1, 65 
NONLINEAR EQUATIONS, 55 
NONLINEAR SYSTEMS, 47 
NOZZLE DESIGN, 54 
NOZZLE FLOW, 54, 62 
NOZZLE GEOMETRY, 54 
NOZZLES, 64 

NUMERICAL ANALYSIS, 14 
NUMERICAL CONTROL, 52 



OGIVES, 11 

ON-LINE SYSTEMS, 47 
OPENINGS, 36 

OPERATIONS RESEARCH, 22 
ORBIT DECAY, 54 
OSCILLATIONS, 12 
OVERPRESSURE, 6 



PACIFIC OCEAN, 22 

PANEL FLUTTER, 67 

PANELS, 67 

PARABOLIC BODIES, 7 

PARACHUTE DESCENT, 25 

PARACHUTES, 37 

PARAGLIDERS, 37 

PARAWINGS, 33 

PAVEMENTS, 59 

PEACETIME, 69 

PERFORMANCE PREDICTION, 41 

PERFORMANCE TESTS, 23 

PERSHING MISSILE, 8 

PILOT INDUCED OSCILLATION, 44 

PILOT PERFORMANCE, 30, 46, 68 

PILOT SELECTION, 30 

PILOT TRAINING, 30 

PILOTLESS AIRCRAFT, 30 

PITCHING MOMENTS, 5, 10, 46, 48 

PLANETARY ATMOSPHERES, 55 

PLANETARY GEOLOGY, 54 

PLANFORMS, 12 

PLATES (STRUCTURAL MEMBERS), 
19 

PLUMES, 58 

POROUS WALLS, 62 

PREDICTION ANALYSIS TECH- 
NIQUES, 42 

PRESSURE DISTRIBUTION, 2, 5, 14, 
38, 53, 60, 63 

PRESSURE DRAG, 14, 16 

PRESSURE GRADIENTS, 70 



PRESSURE MEASUREMENT, 13, 59, 
63 

PRESSURE OSCILLATIONS, 70 

PRESSURE RATIO, 13, 31, 63 

PRESSURE REDUCTION, 5 

PROGRAMMABLE LOGIC DEVICES, 
52 

PROPELLER SLIPSTREAMS, 37, 49, 
50 

PROPULSION, 66 

PROPULSION SYSTEM PER- 
FORMANCE, 57 

PROPULSIVE EFFICIENCY, 66 

PROTUBERANCES, 7 

PROVING, 30 



Q 

QUALITY, 42 



RADAR APPROACH CONTROL, 24 
RADIANT HEATING, 65 
RADIATIVE HEAT TRANSFER, 57 
RAMJET ENGINES, 58, 66 
REAL GASES, 65 
RECEIVERS, 23 
RECESSES, 40 
REENTRY, 37, 49, 53 
REENTRY EFFECTS, 57 
REENTRY TRAJECTORIES, 57 
REENTRY VEHICLES, 36, 41, 49, 55, 

56 
REFLEXES, 68 
RELIABILITY, 32 

REMOTELY PILOTED VEHICLES, 26 
RESEARCH AIRCRAFT, 19 
RESEARCH VEHICLES, 18 
RESONANT FREQUENCIES, 19 
REUSABLE LAUNCH VEHICLES, 40 
REYNOLDS NUMBER, 2, 11 
RINGS, 41 
RISK, 26 

ROBOT CONTROL, 1 
ROBOTICS, 1 
ROCKET THRUST, 63 
ROLL, 35 

ROTARY WING AIRCRAFT, 26 
ROTARY WINGS, 7, 38, 61 
ROTOR AERODYNAMICS, 17, 37 
ROTOR BLADES (TURBOMACHIN- 

ERY), 36 
RUDDERS, 15 
RUNWAYS, 23, 59 



SAFETY MANAGEMENT, 20 

SAMPLING, 32 

SATELLITE ATTITUDE CONTROL, 51 

SATELLITE ROTATION, 51 

SCALE MODELS, 3, 8, 12, 16, 18, 19, 

27, 34, 35, 55 
SCANNERS, 68 
SCHEDULING, 21 
SEAPLANES, 7 
SEMISPAN MODELS, 47 
SEPARATED FLOW, 5, 10, 11, 30, 43 
SHOCK WAVES, 2, 15, 18, 65, 66 
SHORT TAKEOFF AIRCRAFT, 37 
SIDESLIP, 51 
SIMULATION, 24 
SIMULATORS, 38 
SKIN FRICTION, 16 
SLENDER BODIES, 6 
SLOTTED WIND TUNNELS, 53 
SOFT LANDING, 25 
SOFTWARE ENGINEERING, 35 
SOLID PROPELLANT ROCKET 

ENGINES, 57 
SOLID SURFACES, 62 
SONIC BOOMS, 6 
SOUND PRESSURE, 59, 70 

SPACECRAFT CONSTRUCTION 
MATERIALS, 57 

SPACECRAFT CONTROL, 41 

SPACECRAFT LANDING, 36, 55, 56 

SPACECRAFT REENTRY, 36 

SPACECRAFT STABILITY, 41 

SPECIFIC IMPULSE, 54 

SPHERES, 61, 66 

SPILLING, 21 

SPIN DYNAMICS, 3, 16, 18, 25, 51 

SPIN REDUCTION, 51 

SPRAYING, 21 

SR-71 AIRCRAFT, 40 

STABILITY DERIVATIVES, 35, 51 

STABILITY TESTS, 46 

STABILIZERS (FLUID DYNAMICS), 
46 

STAGNATION POINT, 57, 59, 65 

STAGNATION PRESSURE, 58, 63 

STATIC AERODYNAMIC CHAR- 
ACTERISTICS, 44, 61 

STATIC PRESSURE, 5, 13, 58 

STATIC STABILITY, 2, 4, 44, 49, 50 

STATIC TESTS, 19, 61 

STRAKES, 50 

STRUCTURAL ANALYSIS, 19 

STRUCTURAL DESIGN, 33, 60 

STRUCTURAL VIBRATION, 67 



ST-4 



SUBORBITAL FLIGHT, 55 
SUBSONIC FLOW, 18, 52, 70 
SUBSONIC SPEED, 8, 17, 27, 28, 44, 

49, 50, 56 
SUBSONIC WIND TUNNELS, 52 
SUCTION, 61 
SUPERCHARGERS, 64 
SUPERSONIC AIRCRAFT, 43 
SUPERSONIC DRAG, 4 
SUPERSONIC FLIGHT, 6, 7 
SUPERSONIC FLOW, 12, 18 
SUPERSONIC INLETS, 60 
SUPERSONIC JET FLOW, 63 
SUPERSONIC SPEED, 2, 6, 7, 8, 9, 10, 

11, 12, 15, 16, 18, 27, 37, 41, 47, 52, 

60, 61, 66, 67 
SUPERSONIC TRANSPORTS, 38 
SUPERSONIC WIND TUNNELS, 52 
SUPPORT SYSTEMS, 52 
SURFACE ROUGHNESS, 66 
SURFACE ROUGHNESS EFFECTS, 55 
SURFACE TEMPERATURE, 16 
SWEEP ANGLE, 63 
SWEPT WINGS, 3, 6, 7, 9, 11, 15, 27, 28, 

32, 34, 43, 61 
SWEPTBACK WINGS, 4, 5, 9, 11, 14, 

15,19,45,48,51 
SYMMETRICAL BODIES, 65 
SYSTEMS ANALYSIS, 22, 23 



T TAIL SURFACES, 46 
T-56 ENGINE, 21 

TABS (CONTROL SURFACES), 62 
TAIL ASSEMBLIES, 42, 44, 46 
TAKEOFF, 23, 38 
TARGETS, 48 
TECHNOLOGIES, 26 
TEKTITES, 70 
TELEMETRY, 26 
TELEVISION SYSTEMS, 28 
TEMPERATURE, 19 
TEMPERATURE EFFECTS, 57 
TEMPERATURE PROFILES, 58 
TEST CHAMBERS, 60 
TEST RANGES, 26 
TETHERING, 25, 53, 54 
TETHERLINES, 25, 54 
THERMAL DEGRADATION, 61 
THERMAL STRESSES, 67 
THICKNESS, 3 
THIN WINGS, 10 
THRUST, 31, 36, 54, 66 
TILT WING AIRCRAFT, 9 



TIP SPEED, 36 
TTRES, 59 
TOUCHDOWN, 31 
TOWED BODIES, 18 
TOWING, 25 

TRAILING EDGE FLAPS, 9, 11 
TRAJECTORY CONTROL, 47, 53 
TRANSIENT HEATING, 16 
TRANSONIC FLIGHT, 46 
TRANSONIC FLOW, 14 
TRANSONIC FLUTTER, 42, 46, 48 
TRANSONIC SPEED, 5, 6, 7, 8, 10, 13, 

14,16,18,43,45,47,60 
TRANSONIC WIND TUNNELS, 53, 60 
TRANSPORT AIRCRAFT, 11, 24, 29, 

33,39 
TURBINE ENGINES, 24, 39 
TURBINES, 39, 41 
TURBOFANS, 59 
TURBOJET ENGINES, 19, 41, 64 
TURBOSHAFTS, 41 
TURBULENCE, 64 

TURBULENT BOUNDARY LAYER, 32 
TURBULENT FLOW, 15,41 
TWISTED WINGS, 7, 8, 9 
TWO DIMENSIONAL BODIES, 65 



w 



u 



UNSTEADY FLOW, 17 
UNSWEPT WINGS, 9, 10, 11, 18, 51, 66 
UPWASH, 10 

USER MANUALS (COMPUTER PRO- 
GRAMS), 2 
UTILIZATION, 26 



V 

V/STOL AIRCRAFT, 42 
VACUUM CHAMBERS, 52 
VARIABLE SWEEP WINGS, 63 
VELOCITY DISTRIBUTION, 17 
VERTICAL LANDING, 8 
VERTICAL TAKEOFF, 49 
VERTICAL TAKEOFF AIRCRAFT, 9, 

34, 36, 37, 49, 50 
VIBRATION TESTS, 19 
VIBRATIONAL STRESS, 67 
VIDEO DATA, 28 
VIDEO SIGNALS, 28 
VIRTUAL REALITY, 68 
VISUAL PERCEPTION, 68 
VOICE COMMUNICATION, 24 
VORTICITY, 17 
VZ-2 AIRCRAFT, 30 



WALL FLOW, 52 

WALL PRESSURE, 52 

WALL TEMPERATURE, 11, 62 

WALLS, 11 

WATER, 21 

WAVE DRAG, 3, 6 

WEDGES, 16 

WIND DIRECTION, 48 

WIND TUNNEL APPARATUS, 53 

WIND TUNNEL CALIBRATION, 12 

WIND TUNNEL MODELS, 3, 7, 8, 10, 
19, 27, 42, 47, 49, 50 

WIND TUNNEL STABILITY TESTS, 
12,41 

WIND TUNNEL TESTS, 2, 3, 4, 5, 6, 7, 
8,9,10,11,12,14,15,16, 17,18, 
19, 27, 33, 35, 37, 42, 45, 47, 49, 50, 
51,55,56,57,60,61,63,66 

WIND TUNNELS, 12 

WING LOADING, 9, 25 

WING OSCILLATIONS, 51 

WING PANELS, 30, 61 

WING PLANFORMS, 15 

WINGS, 4, 14, 33, 62 



X RAY INSPECTION, 20 
X-15 AIRCRAFT, 4, 13, 27, 31, 33 
X-24 AIRCRAFT, 55 
X-33 REUSABLE LAUNCH VEHICLE, 
40 



YAW, 42 



ST-5 



Personal Author Index 



Abel, Irving, 55 
Adams, E. W., 70 
Adams, J. J., 50 
Alberts, C. J., 1 
Alem, Nabih, 68 
Alexandris, Georgios, 11 
Alksne, Alberta Y., 14 
Allen, Clyde Q., 62 
Ameri, A. A., 40 
Anderson, Roger A., 56 
Anderson, Seth B., 28, 42 
Angelo, Anthony W., 28 
Aoyagi, Kiyoshi, 11 
Annan, Ali, 17 



B 

Babb, C. Donald, 27 
Bamett, D. O., 60 
Bamett, Donald O., 54 
Batterson, Sidney A., 35 
Beheim, Milton A., 60 
Bielat, Ralph P., 45 
Biggerstaff, S., 30 
Blair, A. B., Jr., 27 
Blake, D., 24 
Blower, D. J., 30 
Bohon, Herman L., 67 
Boisseau, Peter C, 48 
Boltz, Frederick W., 62 
Booz, Julieta E., 26 
Bos, A., 2, 19 
Bowers, Albion H., 25 
Bowman, James S., Jr., 3, 12, 17 
Brady, James A., 8, 61 
Brandt, Keith E., 20 
Braslow, Albert L., 5 
Braun, R. D., 53 
Bray, Richard S., 55 
Briggs, Benjamin R., 18 
Browne, James S., 29 
Buchanan, Randy K., 52 
Burbank, Paige B., 59 
Burk, Sanger M., Jr., 3 
Busch, Arthur M., 21 
Buschman, Albert J., Jr., 65 



Campbell, John A., 21 
Capone, Francis J., 17 
Carico, Dean, 25 
Carlson, Harry W., 6 
Carr, Robert E., 66 
Carroll, C. W., 1 
Carter, Howard S., 66 
Cassetti, Marlowe D., 9, 43 



Chapman, A., 30 
Chapman, A. D., 30 
Chapman, Dean R., 54 
Cheatham, Donald C, 56 
Cheatwood, F M., 53 
Church, James D., 10 
Coleman, Thomas L., 32 
Connor, Andrew B., 48 
Conti, Raul J., 16,63 
Corda, Stephen, 39 
Corpening, Griffin P., 39 
Cosmo, Mario L., 53 
Cottrill, Gerald C, 47 
Cox, Timothy H., 39 
Creer, Brent Y., 67 
Cremin, Joseph W., 10 
Croom, Delwin R., 36 
Cubbage, James J., Jr., 52 
Cubbage, James M., Jr., 5 



Daugherty, James C, 2, 14 
Davis, William D., 58 
Decker, John P., 55 
Desai, P. N., 53 
Deveikis, William D., 65 
Dickey, Robert R., 61 
Dixon, Sidney C, 67 
Dougherty, Michael R., 22 
Draper, Mark H., 68 
Drinkwater, Fred J., 55 
Drinkwater, Fred J., IJJ, 44 



Eggleston, John M., 42, 53 
Elkan, Elizabeth, 22 
Engelund, W. C, 53 
Esgar, Jack B., 25 
Estes, Robert D., 53 



Falanga, Ralph A., 63 
Farmer, Moses G., 41 
Fetner, Mary W., 29 
Fichter, Ann B., 9 
Fink, Marvin P., 18 
Fobes, J. L., 20 
Forrette, Robert E., 40 
Foster, Gerald V., 41 
Founder, Roger H., 27 
Fox, Annie G., 46 
Frankenberger, C. E., 41 
Frankenberger, C. E., JJI, 39 
Fukui, Tomomi, 58 



Ganvert, E., 19 
Gates, Ordway B., Jr., 47 
Gault, Donald E., 61 
Gera, Joseph, 25 
Gennann, Kenneth Paul, 23 
Gessow, Alfred, 36 
Gracey, William, 38 
Green, Kendal H., 2, 14 
Greer, Daniel S., 37 
Greiling, Y, 19 
Griffith, George E., 67 
Griner, Roland F, 42 
Groen, Joseph M., 60 
Gronlund, Scott D., 22 
Gruenwald, David L., 69 
Granwald, Kalman J., 37, 49, 50 
Gustafson, F B., 36 



H 

Hall, Albert W., 38 
Hall, Leo P., 61 
Harris, Jack E., 38 
Hassell, James L., Jr., 4, 27 
Hasson, Dennis F, 9 
Hatfield, Elaine W., 2 
Hayashi, Hidechito, 58, 59 
Healy, Frederick M., 3, 12, 17 
Heinle, Donovan R., 45 
Herrmann, M., 54 
Hewes, Donald E., 27, 33 
Hey son, Harry H., 17 
Hickey, David H., 11 
Hicks, John M., 28 
Hill, R., 24 

Holdaway, George H., 12 
Holdaway, George H., 2, 8 
Holeski, Donald E., 40 
Holzhauser, Curt A., 9 
Hopkins, Edward J., 5 
Horan, Colleen, 5 1 
Houbolt, John C, 35 
Howes, Walton L., 70 
Hucke, William L., 34 
Huffaker, R. M., 70 
Hunter, Paul A., 24 
Hurt, George J., Jr., 46 
Huston, Robert J., 37 
Hyett, B. Jeanne, 59 



Igoe, William B., 9 
Innis, Robert C, 28 



PA-1 



James, Carlton S., 30 
Janos, Joseph J., 63 
Jewel, Joseph W., Jr., 24, 29 
Jillie, DonW., 5 
Johnson, Joseph L., Jr., 4, 48 
Johnson, L., 54 
Johnson, Virgil E., Jr., 7 
Jones, George W., Jr., 41 
Jones, Robert A., 57, 65 



K 

Kangas, J. A., 53 
Kaufman, W. M., 1 
Keating, Stephen J., Jr., 10 
Kelley, Henry L., 36 
Kelly, Thomas C, 18 
Kenyon, George C., 62 
Keyes, J. Wayne, 18 
Klinar, Walter J., 25 
Koczo, S., 21 
Kodama, Yoshio, 58, 59 
Koenig, David G., 8 
Kolnick, Joseph J., 59 
Kopardekar, Parimal, 22 
Kraszewski, Edwin T., 66 
Kuhn, Richard E., 37, 50 



Lancashire, Richard B., 58 
Land, Norman S., 46 
Lansing, Donald L., 6 
Larson, Howard K., 10 
Larson, Richard R., 39 
Laurence, James C., 64 
Lazzeroni, Frank A., 2 
Lee, Henry A., 35 
Levin, Alan D., 5 
Levine, Jack, 57 
Levy, Kionel L., Jr., 15 
Levy, Lionel L., Jr., 14, 56 
Lezberg, Erwin A., 58 
Libbey, Charles E., 3, 35 
Lichtenstein, Jacob H., 51 
Lippmann, Garth W., 10 
Little, B. H., Jr., 52 
Lokos, William A., 25 
Lorenzini, Enrico, 53 
Lovelace, Uriel M., 57 
Lovell, J. Calvin, 54 
Loving, Donald L., 4, 13 
Lowiy, John G., 36 



M 

Manning, Carol A., 22 
Margrif, Frank, 64 
Marker, T., 24 
Martin, James A., 55 
Martz, C. William, 57 



Matranga, Gene J., 31 
McCloud, JohnL., EI, 61 
McDuffee, Michael, 64 
McFadden, Norman M., 45 
McGuire, Robert J., 32 
McKay, James M., 33 
McLaughlin, Milton D., 24 
McNeil, Michael, 21 
McShera, John T., 7, 18 
Meadows, May T., 32 
Melillo, Michael R., 51 
Mellenthin, Jack A., 8, 12 
Menon, P. K., 43 
Mertaugh, Lawrence J., 26 
Mills, Scott H., 23 
Moes, Timothy R., 39 
Mogford, Leslye S., 22 
Mogford, Richard H., 22 
Monaghan, Richard C, 39 
Morelli, Eugene A., 26 
Morgan, William C, 25 
Mottard, Elmo J., 66 
Moul, Martin T., 49 
Murahata, Kazuhiro, 58 
Murray, James E., 25 



N 

Neal, Bradford A., 39 
Neely, Robert H., 42 
Neiderman, Eric C, 20 
Neihouse, Anshal I., 25 
Nelson, Richard D., 2 
Normyle, Dennis, 26 
Notarianni, Kathy A., 58 
Nuzman, Edward F, 51 



Ohrt, Daiyl D., 22 
ONeal, Robert L., 15 



Page, Anthony B., 43 
Parlett, Lysle P., 49 
Patterson, Elizabeth W., 5 
Patterson, James C, Jr., 4 
Paulson, John W, 33 
Peeiy, H. Rodney, 47 
Pegg, Robert J., 29, 48 
Peio, Karen J., 51 
Perlee, C. J., 1 
Perry, Jennifer L., 22 
Peters, Todd L., 25 
Peterson, Victor L., 44 
Phillips, William H., 35, 42 
Pippin, Bradley W, 20 
Pittman, Claud M., 65 
Pitts, William C, 13 
Plohr, Henry W., 40 
Polamus, Edward C, 50 
Portman, C. A., 30 



Powers, Brace G., 39 
Press, Harry, 32 
Prinzo, O. V, 23 



Quigley, Hervey C, 28 



Raeth, Peter G., 68 
Rashis, Bernard, 62 
Rasnick, Thomas A., 7 
Re, Richard J., 9 
Reinhardt, J., 24 
Reising, John M., 68 
Reshotko, Eli, 65 
Richard, Michael, 64 
Rigby, David L., 40 
Ritchie, Virgil S., 4 
Rodgers, Mark D., 22 
Rogallo, Francis M., 36 
Roosa, John D., 69 
Rosecrans, Richard, 60 
Ruhlin, Charles L., 46 



Sadoff, Melvin, 44, 45 
Salmi, R. J., 60 
Saltzman, Edwin J., 4, 12 
Samanich, N. E., 60 
Samanich, Nick E., 54 
Sammonds, Robert I., 15 
Sarkos, C, 24 
Sawyer, Richard H., 59 
Scallion, William I., 7 
Scher, Stanley H., 16, 25 
Schmeer, James W., 43 
Schy, Albert A., 49 
Scott, William R., 31 
Seiff, Alvin, 32 
Sevier, John R., Jr., 5 
Shanks, Robert E., 33 
Sharkey, Robert, 21 
Shoemaker, Charles J., 66 
Siegel, M. W., 1 
Sieipien, Larry, 64 
Smedal, Harald A., 67 
Smith, Charles C, Jr., 34 
Smith, Donald W., 59 
Speitel, L. C, 24 
Spreemann, Kenneth P., 13, 50 
Spreiter, JohnR., 14,59 
Stahl, David, 22 
Stallings, Robert L., Jr., 59 
Steinberg, Marc L., 43 
Steinthorsson, E., 40 
Stell, Richard E., 38 
Stickle, Joseph W., 38 
Stienberg, M. L., 43 
Stivers, Louis S., Jr., 15 
Stokes, George M., 52 



PA-2 



Stroud, W. Jefferson, 19 £ 

Swain, Robert L., 57 

Swann, Robert T., 56 Zaman, K. B. M. Q., 61 

Swanson, Andrew G., 57 

Swenson, Byron L., 16 

Sweriduk, G. D., 43 



Tambor, Ronald, 28 
Tanaka, Kiyohiro, 58, 59 
Tapper, Phillip Z., 58 
Tapscott, Robert J., 36 
Taylor, Robert T., 36 
Thome, David K., 64 
Thomson, Robert G., 66 
Townsend, Quwatha S., 7 
Trescot, Charles D., Jr., 10 
Trimpi, Robert L., 65 
Tuovila, W. J., 46 
Turner, Thomas R., 11 



u 



Ulbrich, Norbert, 51 
Unangst, John R., 48 



V 

VanHise, Vernon, 2 
Vicente, James P., 51 
Vidal, A., 19 
Voelker, Leonard S., 39 
Vomaske, Richard R, 44 
Vu, Tong, 32 



w 

Walker, Robert W., 65 
Weiberg, James A., 9 
West, F. E., Jr., 6, 10 
Whitcomb, Richard T., 5 
White, Maurice D., 55 
White, William L., 16 
Whitten, James B., 46 
Wiggins, Lyle E., 13 
Wiley, Alfred N., Jr., 10 
Wiley, Harleth G., 45 
Wilkins, Max E., 32 
Williams, James L., 49, 51 
Wingrove, Rodney C., 67 
Winston, Matthew M., 37 
Witte, William G., 62 
Wong, Norman, 9 
Woodling, C. H., 47 
Wornom, Dewey E., 7 



Yamaguti, Akihiro, 59 
Young, John W., 53 



PA-3 



Report Documentation Page 



1. Report No. 
NASA/SP— 1998-7037/SUPPL387 



2. Government Accession No. 



3. Recipient's Catalog No. 



4. Title and Subtitle 

Aeronautical Engineering 

A Continuing Bibliography (Supplement 387) 



5. Report Date 

November 13, 1998 



6. Performing Organization Code 



7. Author(s) 



8. Performing Organization Report No. 



10. Work Unit No. 



9. Performing Organization Name and Address 

NASA Scientific and Technical Information Program Office 



1 1 . Contract or Grant No. 



12. Sponsoring Agency Name and Address 

National Aeronautics and Space Administration 
Langley Research Center 
Hampton, VA 23681 



1 3. Type of Report and Period Covered 
Special Publication 



14. Sponsoring Agency Code 



15. Supplementary Notes 



16. Abstract 

This report lists reports, articles and other documents recently announced in the NASA STI 
Database. 



1 7. Key Words (Suggested by Author(s)) 
Aeronautical Engineering 
Aeronautics 
Bibliographies 



18. Distribution Statement 

Unclassified - Unlimited 
Subject Category - 01 



19. Security Classif. (of this report) 
Unclassified 



20. Security Classif. (of this page) 
Unclassified 



21 . No. of Pages 
94 



22. Price 

A05/HC 



For sale by the NASA Center for AeroSpace Information, 7121 Standard Drive, Hanover, MD 21076-1320